Liquid rocket space engines

NASA TECHNICAL TRANSLATION NASA TT F- 1750 3

LIQUID ROCKET SPACE ENGINES
V. N. Bychkov, G. A. Nazarov and V. I. Prishchepa

(i»AaA-TX-P- 17503) ^±^Uiu BGLKtx SPACt K77-1J14U

citKiLhSS {Joint EutiicatioiiS aesearch
Service) 5o f bC A^j/mz AOI C5CL 2lti

Uacias

Translation of "Kosmicheskiye zhidkostno-raketnyye dvigateli,"

Novoye v zhizni, nauke, tekhnike; Seriya "Kosmonavtika,

astronomiya , " No. 9, 1976, pp. 1-64

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON, D.C. 20546 DECEMBER I976

I. R«po>t No

__NASA TT F-17503

4. Tifle ond Subfillv

LIQUID R0C3CET SPACE ENGINES

2. Covornment Accvssion No.

7 Aothofi »)

V. N. Bychkov, G. A. Nazarov,
V. I. Prishchepa

SIANOAHO Tin f PACE
3. Rectptent's Cotolog No

S. Ropo'i Ooio

December 1976

6. Pofformtng O'Qoni totion Code

6. Performing Orgoniiolion Report No.

9. Performtng Orgonixotion Nome and Address

Joint Publications Research Service
_ Arlington. VA 22201

Work Unit No

I I Controct or Gront No.

W- 1318 3

12 Sponso'ing Agency Nam« ond Address

National Aeronautics and Space Administration
Washington, D.C. 20546

13. Type of Report ond Period Covered

Translation

14. Sponsoring Agency Code

15 Suppfemen*ary No*es

Translation of "Kbsmicheskiye zhidkostno-raketnyye dvigateli," Novoye
V zhizni, navike, tekhnike; Seriya "Kosmonavtika, astronomiya , " No. 9,
1976, pp. 1-64.

16 Abs"oct

Liquid rocket engines are the main engines of space rockets and space-
craft. One of the outstanding achievements of mankind — launching of
the first artificial earth satellite — was accomplished by using liquid
rocket engines. And somewhat later powerful liquid rocket engines
inserted the first cosmonaut of the planet — Yu. A. Gagarin — into
orbit around the earth. These engines, their design and use in cosmo-
nautics are discussed in the pairphlet. It is intended for a wide range
of readers.

17. Key Words (selected by Author(s))

19. Security ^lossif. (of this report)

Unclassified

18. Distribution Stotement

Unclassified - Unlimited

20. Security Clossif. (of this poge)

Unclassified

2). No. ->f Pogos

50

22. P„c

NASA HQ

Liquid rocket engines (ZhBD) , which operate on liquid fuel, along with solid
fuel rocket engines (RDTT) are the main types of space engines. They are
used extensively in cosmonautics. Man's daring dream — travel into space

— was accontplished by using liquid rocket engines. On 4 October 1957 the
world's first launch of an artificial earth satellite, which opened the
space age in mankind's history, occurred in the USSR using powerful liquid
rocket engines and on 12 October 1961 liquid rocket space engines provided
orbital insertion of a spacecraft euround the earth with the first cosmonaut

— Yu. A. Gagarin.

Almost all space rocket launches are now accomplished using liquid rocket
engines. Liquid rocket engines are used extensively and directly in space
during spacecraft flints: artificial earth satellites, automatic inter-
planetary stations, spacecraft* and so on.

The numerous liquid rocket space engines may be divided into three groi:ps:
the main engines of space rockets, the main engines of spacecraft and
auxiliary engines. Main liquid rocket engines are used in space operations
which require high specific expenditures: during acceleraticni of space
rockets, during acceleration and braking of spacecraft, transfer of space-
craft to other orbits and so on. Auxiliary engines support operations un-
related to high energy expenditures: staging and space rocket flight con-
trol, spacecraft orientation and stabilization ari so on. It should be
said that main liquid rocket engines also frequently fulfill avixiliary func-
tions. Moreover, clear delineation of liquid rocket engines into main and
auxiliary is not always possible.

General Data on the Operating Principle of Liquid Pocket Space Engines

Liquid rocket engines operate on the principle of conversion of the potential
chemical energy of liquid fuel to the kinetic energy of gases flowing from
the engine; in this case an emissive force occurs, directed opposite to the
gas flow and called reactive force or thrust. It should be noted that a
liquid rocket engine is only part of a rocket power plant which also contains
fuel tanks, fittings and pipelines which connect the liquid rocket engine to
the tanks.

*Space rockets are also frequently related to spacecraft in scientific and
technical literature. In this paitphlet we shall call spacecraft those
enumerated objects which are space rocket pay loads.

REPRODUCIBILITY OF liii,
ORIGINAL PAGE IS POOR

Liquid rocket engine fuel may be bxpropellant and monopropellant. Bipro-
pellant consists of a liquid oxidizer (oxygen, nitrous oxides and so on)
and a liquid fuel (hydrogen, hydrocarbons and so on) stored in sep6u:ate
tanks. Mot propellant is a liquid (for example, hydrazine) capable of
catalytic decomposition.

Conversion of liquid fuel to a reactive gas jet occvirs in a chamber which
is che main and indispensable component of any liquid rocket engine. The
chamber operating on bipropellant contains a combustion chamber in which
the oxidizer and the fuel interact with each other (burn) with formation of
a high- temperature gas, and a propelling supersonic nozzle in which the
formed gas is accelerated to a velocity exceeding the speed of sound. Total
fuel ccanbustion is achieved by a preliminary atomizer and by mixing of the
oxidizer and fuel using a mixing head equipped with injectors. The gas
tenperature in the chamber reaches several thousand degrees and continuous
cooling of it is therefore required to maintain the integrity of the chamber
under these conditions. It may be accomplished, for example, by using
fuel flowing through channels in the chamber housing prior to entering the
mixing head. This nethod of oxidation is called regenerative.

A cheimber operating on a bipropellant and its working process are shown in
Figure 1.

2 3 4 5 6 7

Figure 1. Diagram of Liquid Rocket Engine Chamber Operating on
Bipropellant and Variation of Gas Parameters Along
Its Length 1: 1 — oxidizer; 2 — mixing head; 3 —
combustion chamber; 4 — subsonic part of nozzle; 5 —
supersonic part of nozzle; 6 — cooling channel; 7 —
fuel

A liquid rocket engine chamber operating on monopropellant contains a so-
called expansion chamber in which the liquid fuel is converted to a gas in
the presence of a catalyst. It may either by the liquid entering the
chamber from a separate tank or a solid located in the chamber itself.

Liqvdd rocket space engines are divided into pressure fuel feed and with
pun^ feed engines. The fuel in liquid rocket, engines of the first type
enters the chamber from tanks due to being forced out by gsises entering,
for exanple, from special tanks. A liquid rocket engine with pump feed is
distinguished from that described by the presence of a rurbopump unit
and gas generator. The turbopump unit contains fuel pumps (usually of the
centrifugal type) and a gas turbine which sets them in motion, rotated by a
gas vAiich is produced in the mentioned gas generator; spent gas may be ex-
hausted to an exhaust pipe.

A liquid rocket engine with pun^ fuel feed is shown in Figure 2 in the
composition of an engine power plant.

tpMcp)

Figtire 2. Diagram of Engine Power Plant With Liquid Rocket Engine
Having Pump Feed of Fuel: a — liquid rocket engine;
1 — turbine exhaust pipe; 2 — oxidizer pump; 3 —
fuel puirp; 4 — fuel tank; 5 — oxidizer tank; 6 —
turbine; 7 — gas generator; 8 — chamber

In the general case the thrust of a liquid rocket space engine consists of
the chamber thrust (the main component) and the thrust of the turbine ex-
haust pipes. The chamber thrust is determined in this case as P)^ * nit*Ja ■*■
■•■ fa(Pa ' Pn) f where m is the fuel mass expended by the chamber within
1 second; Wg is the gas flow velocity; Fg is the output cross- sectional
area of the chamber; p^ is the gas pressure at the chamber output; and p^
is the pressure of the surrounding medium.

The ratio of thrust to fuel consumption (denoted as lu) has the dimension
of velocity suid is called specific impulse (of a liquid rocket engine or
the chamber). This parameter characterizes liquid rocket engine efficiency.
It obviously follows from the well-known Tsiolkovskiy formula, which may be
written as In Mq/Mjc = vj^/I^ (here Mfe and Mq are the final and initial masses,
respectively, of a rocket unit and vjc is the final velocity of the apparatus) ,
but even a sli^t increase of lu leads to a significant increase of the per-
missible ratio %/Mo, i.e., all things being equal, to a significant in-
crease of the rocket pay load. As can be seen from the given formula for
P]c, the thrust and specific impulse of a liquid rocket engine increase as
flight altitude increases. Since the first term in this formula is consi-
derably greater than the second and, moreover, since the main fuel mass of
the liquid rocket engine flows through the chamber, the specific impulse of
a liquid rocket engine is approximately equal to the gas flow rate from the
chambe r and may be calculated in the following manner: ly ^ Wg =
= 129 1/Tk[l - (pa/Pk)'*' ^/fi^i where Tjc is the gas temperature in the combus-
tion chamber (expansion) ; p]^ and pg are the gas pressure in the combustion
chamber and at the nozzle output, respectively; /* is the molecular weight
of the gas in the combustion chamber; and ^ is a coefficient vrtiich charac-
terizes the thermophysical properties of the gas in the chamber (usually
A ^ 15) .

The values of parameters t^, /u emd A are mainly determined by the type of
rocket fuel. If the specific heat of the fuels is identical, a greater
value of l^ and consequently of lu corresponds to products of greater calori-
fic value. Parameter ^ determines the gas productivity of the fuel: the
smaller jjl , the more gases are formed in the chamber and the higher is Iu>

Much higher values of lu correspond to bipropellants than to monoprope Hants.
Unfortunately the fuels which provide the highest specific inpulse have low
density and consequently require cumbersome (emd heavy) fuel tanks for their
disposition, which reduces the efficiency of these fuels.

If parameters Tj^, y- and A are determined by the type of rocket fuel, the
values of pj^ and Pa are selected during engine design. Liquid rocket
engines which launch space rockets are usually designed for pa = 0.4-0.7 atm
and those operating during the last stages of space flight are usually de-
signed for smaller values of Pa (up to hundredths of an atmosphere) . When
selecting the value of Pa, the fact that the overall dimensions and weight
of the nozzle increase as this parameter decreases is particularly taken
into account.

An increase of pressure in the combustion chamber (p^) is a real means of
increasing the specific impulse since the ratio Pa/Pk decreases as pressure
P]^ increases (for a selected value of pg) .

An increase of pressure p^ also leads to a decrease of the overall dimen-
sions of the chamber. For these reasons the pressure in the chamber is
related to the nuitiber of main parameters which characterize the degree of
technical perfection of liquid rocket engines. The value of p^ for liquid

rocket space engines with pressure fuel feed usually comprises approximately
10 atm; the value of p]( for liquid rocket engines with pump feed is consi-
derably higher and reaches hundreds of atmospheres. The latter engines are
better than those of the first type both in economy and overall dimensions
and in weight ratio: they are characterized by a small specific weight (the
weight of liquid rocket engines in kilograms per 1 ton of thrxist) .

Although the fuel tanks in the case of a liqviid rocket engine with pump feed
are under gas pressure (tank boost pressure) , this pressure is less than in
the case of a liquid rocket engine with pressure feed and consequently the
fuel tanks are lighter.

Puitp feed is now used only in liquid rocket engines operating on bipropel-
lants. With rare exceptions, all llqtdd rocket engines with thrust greater
than 10 tons contain txirbopump assemblies.

Let us note in conclusion that the following are taken into account when
selecting the type of rocket fuel, operating parameters and layout of liquid
rocket engines (along with the general concepts outlines) : the possibility
of storing the rocket fuel under space flight conditions, the operating
mode of the liquid rocket engine, the reliability of the rocket engine
power plant and many other factors.

The History of Development and Use of Liquid Rocket Engines

The liquid rocket engine was first proposed by our own K. E. Tslolkovskly
in 1903 as an engine for space flight. Tslolkovskly defined development of
powerful, economic liquid rocket engines as a primary problem on the path
toward space flight.

Practical investigations on development of liquid rocket engines were begun
in 1921 by the Americem R. Goddard, who somev^iat later in 1926 launched a
small rocket with a liquid rocket engine. Development of liquid rocket
engines was begun in the USSR, Germany and other countries at the end of
the 1920 's and beginning of the 1930' s. The first Soviet experimental
liquid rocket engine, the ORM-1, designed by V. P. Glushko and developed at
the Gas Dynamics Laboratory (GDL) , was tested in 1931.

Experimental models of liquid rocket engines with thrust up to several hun-
dred kilograms, designed for experimental flying vehicles, appeared in the
USSR and the United States prior to the beginning of World War II. Serially-
produced liquid rocket engines were developed in a number of countries by
the end of the war. The first Soviet serially-produced liquid rocket en-
gines were those of the RD-1 type with a thrust of several hundred kilograms,
designed for aircraft. They were developed at the experimental design office,
which was subsequently known as GDL-OKB.

An Important engineering achievement was development of the first large
liquid rocket engines developing a thrust of more than 25 tons in the 1940 's.
Ballistic missiles with a ranc? of several hundred kilometers and also

geophysical rockets which lifted scientific apparatus to high altitudes
were developed on the basis of these liquid rocket engines.

By the middle of the 1950 's liquid rocket engines were subjected to a number
of improvements and the range of the rockets exceeded 1,000 km. The capa-
bilities of the type of liquid rocket engine developed by that time were
essentially exhausted in this case due to the low specific heat of the rocket
fuel (oxygen-ethyl alcohol) used at that time, the inefficiency of its use
and also the imperfection of the engine design of this type

The time that ballistic missiles designed for a range of several thousand
kilometers and of space rockets, development of which was begun in the USSR
and the United States in 1954-1955, appeared was determined largely by the
capability or developing a new type of liquid rocket engine which far ex-
ceeded existing engines in all main characteristics. Investigations to
develop new types of liquid rocket engines, begun in the USSR and the Ur ited
States even during the second half of the 1940 's and which were conducted
parallel with improvement of existing engines , became one of the main pre-
requisites for development of space rockets.

The Soviet space rocket which was used to launch the world's first artificial
satellite in 1957 was equipped with a liquid rocket engine developed at
GDL-OKB under the supervision of the prominent Soviet scientist V. P.
Glushko.

Main Liquid Rocket Space Engines

As noted above, liquid rocket engines are the main types of engines which
are used to accelerate most space rockets. The number of all models of
liquid rocket engines used for this purpose is many thousands. Most Soviet
liquid rocket engines were developed at GDL-OKB; they were installed on all
Soviet rockets launched into space. Development of liquid rocket engines
for space rockets abroad is related primarily to the American Rocketdyne
Company and to the work of a group of engineers supervised by S. Hoffman.
The liquid rocket engines of this company were used to boost more than 80
percent of the American space rockets launched since 1958.

Development of space rocket engines is related to enormous expenditures of
funds. For this reason only four countries except the United States and
the USSR have been able (1965-1971) to launch space objects (artificial
earth satellites) through their own efforts. These countries include
France, Japan, the Chinese People's Republic and Great Britain (enumerated
in the chronology of tlie first launches) .

Having launched only a single artificial earth satellite through their own
efforts, Great Britc.in, like a number of other countries, now enjoys the
use of United States services to launch space apparatus. Japan, being
limited to several launches of low-power space rockets, has now begun to
produce an American rocket under license, on one stage of which a liquid
rocket engine of Japanese design developing a thrust of several tons has

been mounted. France continued to carry out infrequent launches of space
apparatus by using a low-power rocket with a liquid rocket engine on the
first stage. Some West European countries (France, West Germany and others)
are developing the Europeaui "Ariane" space rocket with liquid rocket engine
through their combined efforts. India is developing a space rocket with
RDTT from a model of the American "Scout" rocket. Launch of satellites
using the indicated rockets is planned at the end of the 1970 's and begin-
ning of the igSO* s.

We shall subsequently devote our attention primarily to powerful liquid
rocket space engines with a thrust of several tens and hundreds of tons.
Therefore, the main position in the pamphlet will be devoted to engines of
the USSR and the United States since liquid rocket space engines in these
countries, like cosmonautics as a whole, have been developed extensively.
Let us begin the story of these liquid rocket engines with the world' s first
space rocket engines.

The First Space Rocket Engines (History of Development, Design and Parameters)

The liquid rocket engines designed by GDL-OKB for the first space rocket
were the result cf extensive scientific research and experimental design
work begun in the USSR 10 years before the flight of the first space rocket.
New designs, fuels, materials, production processes and equipment were
developed and assimilated as a result of these investigations. The main
technical solutions utilized in the first space engines were primarily
checked at GDL-CKB during preliminary development of several types of
liquid rocket engines brought to different stages of accomplishment. The
first of them was an engine with a thrust of more than 120 tens, designed
to use oxygen-kerosene bipropellant. Development of this liquid rocket
engine during the period 1947-1951 reached the stage of manufacture and
testing of various subassemblies.

The primary result of all these investigations was development of the liquid
rocket engine chamber, capable of operating at high temperatures and pres-
sures. The chamber design used in the large liquid rocket engines of the
1940' s clearly did not meet this requirement. The chamber walls had to be
made thick from strength concepts and the chamber was massive and heavy.
Strength requirements very soon entered an unresolvable contradiction with
cooling requirements upon augmentation of the charier operation.

A chamber of essentially new design (Figure 3) was suggested, developed and
introduced into Soviet technology. The body of this chamber, designed for
regenerative cooling, consisted of inside and outside walls (called the
fire wall and jacket, respectively) , joined to each other by soldering
through the intermediate corrugated wall (interlining) or through ribs cut
in the fire wall. The chamber contained a mixing head with injectors
soldered into it and was divided technologically into several subassemblies
joined in the final phase of the production process by annular weld seams.

Figure 3. Design of SoMered-WeWed Chaaiser (the Mixing Head of
the ChaAer Is Shown Msoi?e in Cross-Section)

KEY:

1. Oxidizer

2. Fuel

3. Variant I

4. Variant II

Due to the presence of special couplings in the body of this chamber, its
fire wall can be made thin and can be manufactured from relatively weak,
but high heat-condacting copper alloys, • reliable cooling of the chamber is
thus provided. The jacket which receives gas presstire was manufactured frcm
alloy steels providing high strength of the chamber with low weight. The
capability of significantly increasing the economy of liquid rocket engines
appeared with deifelopitient of soldered-welded chambers (due to increasing
the combustion chaiitier pressure and using efficient rocket fuels} with a
considerable reduction of the specific weight of the engine.

Ki*yRODUCiBUJ:ry or 'ii.:.j

♦ jlittllMAL PAGE IB m)?

Numerous test- firings of small experimental chambers preceded manufacture
of full-scale soldered-welded chambers, fls a result the efficiency of the
new type of chamber was checked extensively and the correctness of selection
of oxycjen and kerosene as the fuel was confirmed. This fuel pair exceeded
the oxygen-ethyl alcohol fuel pair then used in liqxiid rocket engines in
both specific impulse and in density and provided the longest flight range
of the rockets . The soldered-welded chambers were designed for a pressure
two- four times greater than that at which oxygen-alcohol liquid rocket
engines operated. The use of kerosene and an increase of pressure resulted
in an increase of teirperature by SOO'C in the ::oinbustion chamber. Subse-
quent development of liquid rocket engines showed, however, that the sol-
dered-welded chamber is also (Efficient vinder much more severe conditions.

Development of soldered-welded chambers largely determined the success of
developing the engines which provided the flight of the first space rocket,
which is a two-stage vehicle consisting of a central pod (second stage) and
four lateral pods (the first stage) . An RD-108 liquid rocket engine was
installed in the central pod and an RD-107 was installed in the lateral
pods (RD is an abbreviation of the name "rocket engine") . All liquid rocket
engines were fired simultaneously, providing a launch thrust of approxi-
mately 410 tons. The lateral pods were separated about 120 seconds after
launch (flight altitude in this case was approximately 50 km and velocity
was 3,200 m/s); the central pod continued to operate for an additional 180
seconds, accelerating the payload to orbital velocity.

The RD-107 and RD-108 engines are similar in characteristics since en
attempt was made to standardize them to the maximum. The main engine
assemblies are distinguished only by the operating parameters (they are
somewhat higher for the RD-107. An overall view and a schematic diagram
of one of the liquid rocket engines (the RD-107) are presented in Figures
4 and 5.

The RD-107 and RD-108 develop a thrust of 102 and 96 tons, respectively,
which is approximately 90 percent developed in each liquid rocket engine
by four identical chambers fed from a single turbopump assembly. A total
of 52 kg of oxygen and 21 kg of kerosene enters the RD-107 chamber within
1 second. The oxygen is delivered directly to the mixing head through a
central pipe and the kerosene is delivered to an annular collector located
near the chamber outlet, from which it is distributed through the cooling
channels and then, heated to 210''C, is fed to the mixing head. The fuel is
vaporized by 337 injectors, of which one is installed in the center and the
remaining ones are located in 10 concentric circles. The kerosene injectors,
v*iich create a protective gas-liquid film near the chamber fire wall, are
installed near the periphery. The heat flux to the wall reaches a maximum
value — more than 14 million kcal* (m^hr) ~1 — in the region of minimum
chamber cross-section. The cooling chamber of the liquid rocket engine is
designed here for maximum coolant flow rate — up to 20 m/s and the fire
wall temperature reaches a maximum value — 380°C. Upon combustion of fuel,
a gas with pressure of 60 atm and temperature of 3250''C forms in the

Figure 4. jRD~107 Liquid Rocket Eagines 1 — - steering chambers;
2 — steering chaisber rotation aBsmmblj; 3 —> pipe-
lines for delivery of OKldizer to steering chambers;
4 — mock-up brackets (not included in the liquid
rocket engine design)? S — main chaiabers; 6 — fraiae
for attaching liquid rocket engine to rocketj 7 —
gas generator? S — turbine? 9 — oxidizer pompj 10 —
fuel p«ip? 11 — chasfoer pressure sensor? 12 — aain
oxidizer valvBr 13 — pipelines for delivery of oxi-
dizer to chambers J 14 — main fuel valve? 15 — pipe-
lines for deliveri*- of fuel to chautoers

coBbustion chaiffcer. Ibe pressure drops to 0,4 ata and temperature drops to
1690*C after the gas passes through the nozzle, fhe gas is accelerated in
this case to a i«locitf of 2,950 m/s, imparting a thrust of ssore than 23
tons to the chamber.

Success in developing the ia3~107 and RD-108 engines largely determined
deveiopnent of a canpact, lightweight turbopump assartbly capable of opera~
ting at sufficiently low fuel pressure in the rocket tanks (Figure 6} . The
turbopuBip asseniJly of these liquid rocket engines contains two main and two
auxiliary pumps and the turbine which rotates thea^ whose output is 5,200 hp
for the SB- 107. Ihe ptmps are of the centrifugal type and single-stage? the
turbine is axial and two-stage. The main punps are used to feed fuel to the
chsiftibers. They are installed coaxially with the turbine and have identical
rotational speed with it — 8,300 rpts.* Ihe puufss are designed to feed

*This value and also those pump and gas generator paraireters given later
refer to the ED- 107 engine.

10

ffi&f

M

Figure 5. Schematic Dia^aia of RD~107 Liquid Itocket Engine

oxfgen at a rate of 226 kg/s mder a pressure of 80 atan and to feed kerosene
at a rate of 91 kg/s under pressure of 95 atm. Cavitationless operation of
the puBf)s is provided by installation of low-pressure axial vanes in front

of "the" main" rotors.

Figure 6. Turbopuffip Asserably of the m-107 and RD--108 Engines;

1 --~ fuel pump I 2 — oxidizer pump; 3 — turbine? 4 —
heat exchanger

The auxiliary ptaps are driven by a boos^* r. toe of the pumps delivers
liquid nitrogen to a vaporizer bidlt into the turbine exhaust collector and
used to deliver the gases for supercharging the rocket fuel tanks. Tlie
other auxiliary pump delivers low-water (82 percent) hydrogen peroxide to
the gas generator, wh-.ch produces vapor-gas to drive the turbine.

11

UODUCiBlLiiV OF
•KIGINAL PAGE IS ?0V

The gas generator of the liquid rocket engine for the first space rocket
was a cylindrical tank vrtiich contained a granular stibstance — a catalyst.
UJ)on passing through it, the hydrogen peroxide was decoiif>osed into a mix-
ture of war.er vapor and gaseous oxygen with a pressure of 55 atm and tem-
perature of 560 "C; this mixture was fed at a rate of approximately 9 kg/s
to the turbine blades. The spent gas with a pressure of 1.5 atm and tem-
perature of 20O'*C was ejected through the exhaust pipes at a velocity of
450 m/s. A thrust of approximately 700 kg was developed in this case.

Development of the RD-107 and RD-108 engines was combined with solution not
only of purely "engine" problems but of problems of general improvement of
rockets which entrusted a number of functions previously not inherent to
them to liquid rocket engines. One of these functions was to provide the
rocket flight with a strictly given trajectory and to control the rocket
position. Rotary steering chambers equipped with hollow journals through
which fuel was fed from the main pumps of the turbopump unit and vrfiich pro-
vided declination of the chambers at an angle of +45°, accomplished by
means of hydraulic drives, were provided in the liquid rocket engine design
to fulfill the indicated function. There are two steering chambers in the
RD-IO? and four in the RD-108 engine. The steering chamber is similar to
the mair chamber in design. However, operating at almost the same pressure
as the main chamber of the RD-107, the steering chamber developed six times
less dirust.

Assemblies to control the thrust and the ratio cf fuel component flow rates
were provided in the engine of the first space rocKet, which permitted
flight support at a previously calculated optimum velocity cuid made it
possible to achieve total synchronous exhaustion of the fuel components
from the rocket tanks.

The oxygen-kerosene fuel used in the KD-107 and RD-108 engines is not self-
combustible; it is ignited by means of pyrotechnic devices introduced into
the liquid rocket engine chamber from the nozzles and which are triggered
when an electric current is supplied. The operation of these liqtiid rocket
engines is controlled by electro-, pneumo- and pyroautomatics .

With regard to all the elements in the set, the weight of tlie RD-107 engine
is equal to 1,155 kg, which corresponds to a specific weight of 11.3 kg/t
(the RD-108 is 95 kg heavier). The specific impulse of the RD-107 and
RD-108 is 3,080 and 3,090 m/s, respectively. For comparison let vis point
out that the best models of oxygen-alcohol liquid rocket engines, which pre-
ceded the appearance of space engines, were characterized by a specific
weight of 17 kg/t and specific impulse of 2,430 m/s.

Brief Simmiary of fpace Rocket Development

A two-stage space rocket with RD-107 and RD-108 engines was used in 1957-
1958 to launch the first artificial earth satellites. A third stage with
a single-chamber oxygen-kerosene liquid rocket engine having thrust of 5 t,
developed at the 0KB headed by S. A. Kosberg, was developed for the

12

indicated rocket to launch automatic interplanetairy stations (AHS) to tJie
moon (1959) and the manned orbital flights of the "Vostok" spacecraft (1961) .
The thiree-stage rocket, designed in this manner, was named the "Vofetok."
Later modifications of it with a four-chamber liquid rocket engine having
a thrust of 30 t in the third stage has been used since 1964 to launch the
"Voskhod" spacecraft and since 1967 to launch the "Soyuz" spacecraft.* AMS,
which were inserted into a circumlunar orbit and which made a soft landing
on the Itoon (1966) , Venus (1970) and Mars (1971) , were launched by four-
stage modifications of the "Vostok" rocket.

"Vostok" carrier rockets provided the USSR for a long time with the leading
role in space research and continue to be used extensively up to the pre-
sent time. Up until 1964 they considerably exceeded in capacity all other
space rockets, amcsig \idiich may be named such liquid- fueled rockets as the
Soviet "Kosmos" (used since 1962) and the American "Thor-Delta" and "Atlas-
Agena" rockets (used since 1960). These American space rockets, like the
"Vostok," were developed by installing additional stages on the just deve-
loped liquid-fueled ballistic missiles with a range of several thousand
kilometers (the first space rocket was a variant of an intercontinental
ballistic missile) .

Itore powerful space rockets than "Vostok" rockets were developed in subse-
quent years in the USSR and the United States. They include the "Proton"
(USSR), "Satum-1" (two versions), "Satum-5" and the liquid-fueled rockets
of the "Titan-3" family with launch RDTT. Space rockets of the "Titan-3"
family were developed on the basis of intercontinental ballistic missiles
and the remaining ones are completely original designs.

The "Proton" and the latest versions of the "Sature-1" and "Titan-3" rockets
have approximately three times the lifting power of the modified models of
the "Vostok" rocket. Many outstanding achievements of cosmonautics, which
will be discussed further during description of corresponding rocket engines,
are related to their use.

The "Satum-5" space rocket exceeds all other rockets in lifting power.
It was developed within the "Apollo" program, whose purpose included deli-
very and landing of astronauts on the moon. A rocket with a launch mass
of approximately 3,000 t was required for this. After six expeditions to
the moon (1969-1972) , the "Satum-5" has been used only once — to launch
the manned orbital station "Skylab" (1973) with a mass of approximately 80 t.

The appearance of more powerful space rockets than those of the "Vostok"
type became possible only due to rem'irkable new achievements in the field
of liquid rocket engine development. They include development of engines
with chambers having thrust of more than 100 t, development of new rocket
fuels and achieving pressures exceeding 100 atm in the combustion chambers.
Vfe shall consider all these achievements later using specific liquid rocket

♦These modifications of the "Vostok" rocket are sometimes called the
"Voskhod" and "Soyuz" by analogy with the names of the spacecraft.

13

engine models as an example and we shall begin ovir story with description
of the F-1 engine vrfiich is used on the first stage of the "Satum-5" rocket.

The F-1 Liquid Rocket Engine

Ihis engine, designed by the Rocketdyne Company, is shown in Figure 7. It
operates on oxygen-kerosene fuel and develops a thrust from 690 (during
launch) to 790 t (prior to stage separation) . The F-1 is a single-chamber
liquid rocket engine with pump feed of the fuel. The gas for driving the
tiirbopun^ assembly is produced in a gas generator by combustion of the main
fuel of the liquid rocket engine (i.e., the fuel on which the chamber
operates) , but with a large excess of fuel (a diagram of this liquid rocket
engine is shown in Figure 2) . ;^proximately 3 percent of the total flow
rate of fuel throu^ the liquid rocket engine is expended through the gas
generator. Gas with a tenjserature of approximately 800<»C, vrtiich is ejected
into the nozzle section of the chamber after driving the turbine, is formed
in the gas generator. The gas generation system of the described type is
used extensively in modern liquid rocket engines (the spent turbine gas is
frequently injected into the exhavist pipe) .

The F-1 exceeds all other foreign liquid rocket engines in chamber pressure,
which coirprises 70 atm. A so-called tubular chamber vrtiose body is formed
by shaped tubes joined to each other by soldering (the tubes form the re-
generative cooling channel of the chamber), is used in the F-1 engine. The
mixing head of the chamber is detachable with openings which perform the
role of oxidizer and fuel atomizer nozzles. Tubular chambers are typical
to foreign liquid rocket engines. One of them is shown in Figure 8.

The F-1 engine is fired by "gravity flow" of the fuel: when the starting
valves are opened, the oxidizer and fuel are delivered to the chamber and
gas generator at tank supercharging pressure in the nonoperating pun^s;
when the fuel in the gas generator is ignited, the produced gas turns the
turbopvm^s assembly. This lavinch scheme, which does not require auxiliary
devices to turn the turbopvunp assembly, was first accomplished in the Soviet
liquid rocket space engine "Proton."

The fuel in the chamber is ignited during launch of the F-1 engine by a
starting fuel (a mixture of triethyl aluminim and triethyl borane) , distri-
buted in a special sleeve and ignited upon contact with the oxidizer. This
method of ignition, called the chemical method, is used extensively in
modern liquid rocket space engines (the launch fuel may be located directly
in the pipelines of the liquid rocket engine) .

The assemblies which control operation of the F-1 are triggered by electric
commands and the fuel pressure selected from the high-pressure mains, which
eliminates the requirement for auxiliary conpressed-gas systems.

Five F-1 engines are installed on the "Saturn-5" rocket: the central engine
is fixed and the peripheral engines are in gimbal suspensions which provide
rotation of the engine in its own plane parallel to its axis to control the

14

Figure 7. F~l Liquid Rocket Engines 1 — gimbal suspension;
2 — oxidizer to chamber feed collector ; 3 — usain
oxidizer valve; < — main fuel valve? 5 — high-
pressure oxidizer pipeline; 6 — high-pressure fuel
pipeline; 7 —» gas generator; 8 — tubular part of
chainberj 9 ■ — reirovable part of nozzle; 10 — assem-
bled collector of spent turbine gas; 11 — heat ex-
changer; 12 •— turbine; 13 ■ — fuel pimp; 14 — oxidi-
zer pump

flight direction and position of the rocket. Let us present son« of its
parameters for clear representation of the scale of the F-1 engine. The
engine mass is 8,400 kg and height is 5.6 s; the chaitiber diameter in the
output section is equal to 3.6 m. The engine consumes approximately 2.5 t
of fuel within 1 second, which is delivered by a turbopump assembly having
capacity of 55,000 hp.

15

Figiire 8. Tubular Chanijer of LiquM Rocket Engine: I — oxidizer
feed pipe; 2 — tnel feed collector? 3 — mining head;
4 — giinbal suspension! 5 •-~ ignition device; 6 — fuel
feei pipe,- 7 — tiisular body of chamber

Hydrogen- Qjcygen Engines

Kerosene -oxygen fuel on which the RD-IO?, KD-IOS aM F-1 engines described
above operate, is used extensively in space rockets. Th# velocity of the
Jet blast reached and exceeded 3,000 m/s with development of this fuel.
Developnerit of liquid rocket engines in the middle of the 1960* s, which
operate on oxygen-hydrogen fuel lAiich exceeds approximately 30 percent the
specific iaptilse of oxygen- kerosene fuel, was of great isfiortance for
further developsent of cosM»nautics.

Althotigh oxygen-hydrogen fwel was proposed as early as 1903 by Tsiolkovskiy,
it did not find application for a long time for reasons related to the
specific properties of hydrogen. &s is known, liquid hydrogen is 14 times
lighter than water and boils at 20«K.* Mixtures of hydrogen «id air are
extrenely flanaable and explosive. For example, the energy of the electro-
static discharge which we sometiiBes feel upon touching a door handle is
tens and hundreds of times greater than the energy required to ignite a
hydrogen-air mixture, in this regard production of inexpensive liquid
hydrogen in large quantities was problematical, design and operation of
liquid hydrogen systeras were related to complex engineering problems and
the txMil tanks for liquid hydrogen were too heavy.

*Let us point out for canparison that liquid oxygen is 14 percent heavier
than water and boils at 90 *K.

16

OxygeB-hydrofen fuel is um m<&d. on the upper stages of space rockats» where
It produces the greatest effect, to exiK^le of this is the ani-wraal
•Centawr" stage med on space rockets of the "Atlas" and "Titaa-3" families
and also the secoM and third stages of the "Sattim-S" rocket, fne fael
tacks of th«e rocket stages, designed to distribute liqaid hydrogen, are
gigantic theraos bottles whose seta! walls are covered with heat- insulating
ixslytaer materials. The heat insulation used in the tmks of the "Sattim-S"
rocket is showj in Figure § as an exasfde. Channels through «*iich gaseous
helii« is deli'B'ered during latmching of the rocket to removB explosive
vapors £rc» the insalation ^ich may be accuaalated there, are prorided in
this insulation.

Figure 9. Wall of Liquid Bydrogen Fuel Tank (the Second Stage of
the "Satum-5" Rocket) : 1 — pcwer sheathing falwiinum
alloy) J 2 and 5 — adhesive layer? 3 — channels for
passage of helius} 4 — • honeyconij structure Cpolyure-
thane materiall i 6 — nyloa-idienol layer? 7 ~~ airtight
coating {synthetic tedlar aaterlall

The heat insulation adds to weight of the oxygen -hydrogen stages. Since
the oxygen-hydrogen fuel is three tines lighter than oxygen-kerosene fuel,
it requires three times acre volume for its distribution with the same
mass. Ihe total weight of the rocket stage design required for 1 kg of
fuel is 40 percent greater for oxygen-hydrogen fuel than for oxygen-kerosene
fael. This disadvantage is ceiapensated with an excess of high specific im-
pulse of oxygen-hydrogen engines. With equal launch aass, a space rocket
operating on oscygen-hydrogert fuel is capable of inserting three tliiies more
payload into orbit than a rocket operating on oxygen-kerosene fuel. The
use of this fuel In the upper stages of the "Satum-S" rocket permits low
circular geocentric orbital injection of up to 140 t and injection onto a
lunar flight trajectory of up to 48.5 t of payload.

Along with high efficiency, oxygen-hydrogen fuels have a nuiaber of other
advantages, aaong which should be noted low combustion temperature {200<*C
lower than for oxygen- kerosene fuel) and nontoxicity of both the fuel it-
self and its combustion products C%*ich are a mixture of water vapor and
gaseous hydrogen! ,

17

-HODUCIBILITY OF TL
.iiGiNM. PAGE IS POOR

Let us now dwell on the KLlO and J-2 engines used in the "Centaar" stage and
on the upper stages of the "Satum~5" rocket, respectively,

The RLlO engine, developed by the Pratt-«hitney Company, develops a thrust
of 6.8 t. It is c±>vious from the schematic diagram of the engine shown in
Pigtire 10 that it is a single-chastoer liquid rocket engine with pisnp feed
of fuel. However, unlike other liquid rocket engines with tuAopump assem-
blies, there is no gas generator in the RLIO: the turbine is rotated by
the gas produced upon heating of liquid hydrogen in the cooling channel of
the chaBfcer, The gaseous hydrogen temperature is only ~70*C? it enters the
chanJber after the turbine, where it burns with liquid oxygen at a pressure
of approximately 28 atm {the mixtxire is ignited by an electric spark plug) ,
The design temperature during the initial nioinent is sufficient to convert
the hydrogen to a gas, which provides spinning of the turbine. The sifpple
schematic diagram of the RLIO engine is explained by the exceptionally high
theriKjdynamic characteristics of hydrogen

Figure 10. Schematic Diagram of KLlO Engine: I — fuel puiro?
2 — fuel lines; 3 — oxidizer pump; 4 — turbine;
5 — chansber

The RLlO is the best liquid rocket engine in specific impulse, which is
eaual to 4,360 m/s. Two of these engines arc installed on gimbal suspensions
on the "Centaur" stage. This stage is used extensively tc launch automatic
interplanetary stations. In 1972 a "Mlas" rocket with a "Centaur" stage
imparted solar escape velocity to the spacecraft (the "Pioneer- 10" automatic
interplanetary station for investigation of Jupiter). Beginning in 1965,
more than 100 RLlO engines have been used in space flights withotjt a single
failure.

"Uie J-2 oxygen-hydrogen engine, developed by the Rocketdyne Companv, develops
a thrust of 104 t. It is single-chairfcer with piartp feed of fuel and has the
characteristic feature that separate turbopum? assetiblies, each of which
consists of a pump and turbine, are provided for the oxidizer and fuel. H
single gas generator to which approximately 2 percent of the fuel expended
through the engine is delivered, is provided for the two assemblies. The
produced gas sets the two turbines into rotatitr. <^<^quentially, after which
it is ejected into the nozzle of the tubular chamber through slots between
the tubes throuah which the fuel flows. The liquid rocket engine is fired
and cut off (as In the KLlO) by valves controlled by gaseous helium. The
turbopump assemJslies are turned during launchina by gaseous hydroaen delivered
from a special tank.

IB

The J--2 engine develops a specific imptilse of 4,170 b/s at a combustion
chasfcer pressure of 55 atm. fhe liquid rocket engine, weighing 1,600 kg,
is attached secured to the rocket or to the gimbal suspension. Five of
these engines are installed on the second stage of the "Saturn-5" rocket
and one is installed on the third stage.

Ldqiiid Rocket Engines Operating on High-Boiling Fuel

Along with oxygen-kerosene and oxygen-hydrogen fuels containing liquefied
gases and called cryogenic fuels, high-boiling fuels whose coiii»nents are
liquids under ordinary external conditicms, are used extensively in space
rockets. "These fuels are inferior to the oxygen-kerosene pair in specific
impulse, but exceed it in density. Ute basis for the use of high-boiling
fuels on large scales in rocket technology was laid by development of long-
range ballistic missiles operating on these fuels at the end of the 1950 's
and beginning of the 1960's. fhe problems of operating the rockets and the
COTjpatibility of fuels with the design materials, related to toxicity and
chemical aggressiveness of hic#i-boiling fuels, were resolved successfully.

One of the first liquid rocket space engines SBploying high-boiling fuels
was the RD-214 engine, used since 1962 in the first stage of the "Kosmos"
series of rockets. The i?D-214 develops a thrust of 74 t and operates on a
nitric acid oxidizer (AK) — a mixture of nitrous oxides and nitric acid —
and the by-product of kerosene conversion. Modern high-boiling fuels in
which the oxidizer is nitrogen tetroxide (AT) and the fuel is asyiroetrical
dimethyl hydrazine (NDH3) or a mixture of it with hydrazine called aerozine
(AK-NDKG fuel occupies an intermediate position), are more efficient. The
disadvantages of high-boiling fuels containing NDMG and aerozine are also
their self-costoustibility upon contact of the fuel ccwsponents. The presence
of these fuels permits developnent of liquid rocket engines simple in design
and reliable in operation with a low specific weight.

Engines which operate on M-NDMG fuel include the single-chamber liquid
rocket engine with thrust of 7.3 t of the Mierican Bell Company, used since
1959 on the universal high-altitude "Agena" rocket stage, and the RD-216
engine, developed by CZSL-OKB in 1958-1960 and used on the first stage of one
of the rockets of the "Kosroos" series. "Hie RD-216 consists of two identical
liquid rocket engines having a ccmnon launch system. Liquid rocket engines
are two-chajrfber with turbc-putnp assefrtolies located between the chambers in
the region of their minimum cross-section. Each liquid rocket engine deve-
lops a thrust of 88 t and has a specific impulse of 2,860 m/s at chamber
pressure of 74 atm.

Among the numerous liquid rocket space engines operating on nitrogen tetroxide
fuels, let us indicate here onlv the powerful engines of the fttnerican Aerojet
Company, used in the first and second stages of the "Titan-3" rocket (in the
second and third stages, respectively, in the variant of the rocket with
launch RDTT) .

The first stage engine of the "Titan-3" space rocket consists of two single-
chai*er LWl-M-S engines which develop a total thrust of 215 t and specific

19

impulse of approxiraately 2,810 a/s at chaiaber pressure of 55 atm. 1t»is
engine is t±ie best aiwsng foreign liquid rocket engines in specific weight —
6.9 kg/t. A single~chainber LR91~AJ-9 engine with thrust of approxiisately
45 t and specific impulse of 3,040 m/s, operating at presstire of 58 atm,
is installed in the second stage of the "fitan-S." The spent turbine gases
in this engine are exhausted into a rotary nozzle, providing control of the
rocket position with respect to the longitudinal axis. When the considered
engines are started, the turbopuntp assenfclies are turned by the powder gases
froBi the pyrostarters installed on the turbine housings.

The described liquid rocket engines of the Serojet Company were used speci-
fically in 1965-1966 on the modified "Titan~2" intercontinental ballistic
missile, which launched the 10 "Gemini" spacecraft.

The Engines of the "Kosroc" and "Proton" Space Rockets

Developing ever more efficient rocket fuels, the designers of liquid rocket
engines at the sam.e time attempted to utilize more fully the chemical
energy contained in the fuels, i.e., to convert this energy with the highest
possible efficiency to the kinetic energy of the engine jet.

The greatest advances in increasing the efficiency of liquid rocket engines
were achieved in the USSR. Pressures several times greater than those for
the RD-107 and RD-108 became possible in the soldered-welded design chajribers
of OJL-OKB, now used everywhere in Soviet liquid rocket engines. The new
inodels of Soviet liquid rocket space engines were designed for ever higher
chajiber pressure. Liquid rocket engines operating at chamber pressure of
75-80 atm appeared 5 years after development of the first space engines.
Among them were the RD-216 and RI>-119 engines. Data on the RD-216 were
given above and now let us devote a few lines to the RD-119 engine.

The R0-119 (Figure 115 has been used since 1962 in the second stage of one
of the rockets of the "Kosmos" series. This engine operates on the oxygen-
asyirmetrical dimethyl hydrazine fuel pair and develops a thrust of approxi-
mately 11 t. The M5-119 is inferior in specific impulse (3,450 m/s) only
to oxygen-hydrogen engines. The hlqh specific ii!f>ulse of the engine is pro-
duced by an 'Efficient fuel, high chamber pressure (80 atm), improved mixing
of the fuel prior to ignition and high degree of gas expansion in the
chamber (to a pressure of approKxmately 0.06 atm).

The M>-119 engine is a single-chamber with pump feed of the fuel. The gas
for driving the turbines is produced by thermal expansion of the fuel in
the gas generator; the temperature required for the beginnlna of expansion
is provid'^d by coinbustiwn of a powder charge located in the gcs generator
(this charge is also used for initial turning of the turbopump asserrfcly) .
The spent turbine gas flows throuoh fixed steering nozzles equipped with
gas distributors, prcvidinct control o? the flight direction of the rocket
stage and Its position. The rotational speed of the turbine and the fuel
pum{.;,i is 21,000 rpm. Titardum anu aluminuir alloys are used extensively in
the design of liqtjid rocket engines; because of this tie PJ>~119 is related

20

by specific weight (IS *6 kg/tl to the best ligaM rocket s^ce nnqiM&m iijitii
thirost of several tons.

Figtir© 11. H)-119 LlquM Rocket Ea.gi«e

ffee chaufeer pressmres of the SD~119 and BD-216 are close to the aaxiMiro
pXBSBum fox liquid, t^ocket: engines t in wMch the Bpen^
eje'-ted into the esiiaust pipes or into the chamber nozzle, the fact is
that the design strength of the turbine which rotates the fuel pxmps liasits
the gas temperature In the gas generator to approximate If 800 *C. Both the
temperatttre aM presstire in all the previous ly B«iitioned engines are reduced
significantly after the fas passes through the turbiae (for exainple, tem-
perature is eq»l to -6e0-*e -at px-essore of 1-25, atm in the ^^®^

As we already know, gas with such lew paraneters may not prodace a high
specific iispalse. Thus, although the portion of the liquid rocket engine
fuel yiich is expended to drive the turbopurap asseiibiy does develop thrast,
the specific impulse of the engine Is below that ^ich wottM be pxod-uced If
all the fuel were igMted in the coit>Mtion:chairfc^^^ For example, the
specific iiffttjlse for the H3-216 engine Is 1-1.5 percent below the specif ic
istpalse of the chamber.

fhe head of the fuel pianps and consequentlf their output must be increased
to increase the chaiit>er pressure, f o do this, the flow rate of the gas
rotating the turM-neiu^t fhuu, the fraction of engine

ftiel expended through the gas generator increases'a's the chamber pressure
increases, which leads to sloping of the Increase of specific Iropulse for
the liquid rocket engine of the described scheme and also to a decrease of
this parameter.

21

Figure 12. Diagraa of Liqaid Rocket Engine With Afterburnings
1 — first stage of fuel puaip? 2 — second stage of
f-j6l wmot 3 — OJcMlEer niannj 4 — ■ tiirbln«.s S ~-

f-jel pmapf 3
gas general :ori

— oxidizer panpi
6 — chamber I 7

4 — turbine;
— fuel; 8 —

oxidizer

A significaat increase of chaiftjer pressure in liqaid rocket enfines develop«a
after the M>-216 and SD-llS engines was achieved by developing an engine
scteae with afterburning CFigtare 12) . The greater portion of all the con-
soaed fuel enters tbe gas generator in these engines, Ihe gas pressure in
the gas generator is appro%ii«ately doable that of the gas in the combustion
chaj*er. fhe spent turbine gas enters the chamber stiKing head for after-
burning with the remaining part of the fuel at a pressure of 150 ata or more*

The indicated engines, characterized bf high specific ispulse and efficiency,
have been used extensively since the middle of the I960»s in Soviet rocket
technology. Specif icallf, they have; been used since 1965 on all stages of
the "Proton" space rocket. Sei«ral H)-2S3 engines designed by C»D]>OKB, which
significantly exceed the thrust of the M>--107 and M)~108 engines, were in-
stalled on the first stage of this rocket. So»e data on the ro-253 engine
are presented below.

fhe RD-253 is a single-chamber engine operating on high-boiling calorific
self-igniting fuel. The major portion of the fuel after the pua^ss {approxi-
mately 75 percent} enters the gas generator welded to the turbine housing
and the remaining part is delivered to the regenerative cooling dtannel of
the chafflber. The gas after the turbine is delivered through a gas line to
the coifcustion chairber, where it is burned with the liquid component that
has passed through the chamber cooling channel.

Ihe combustion chafflser pressure of the RD-253 is approximately three times
greater than the corresponding pressure for the IP- 108. For reliable cooling
of the chamber, its fire wall is protected by a refractory ceraaic coating

22

and qas-liquid film formed by delivery of the fuel component from the
cooling channel to the wall through openings in it. The total power of
the turbopujip assemblies of all the ia>-253 engines contained in the propul-
sion plant exceeds 150,000 hpj a power of 100 hp is concentrated in a sin^'le
kilograffi of weight of the turbopump assembly (this assembly operates at a
speed of 14,000 rpm) . The maximum fuel pressure in the main lines of the
engine reaches 400 atm. the design reliability of the engine at such high
pressure is provided by extensive use of welding to join the assemblies and
subassCTiblies.

The liquid rocket engine is fired by "gravity flow" of the fuel (similar to
the P-1 engine) without the use of special starting devices. The use of
self-igniting fuel eliminates the need for an ignition system. The liquid
rocket engine is fired and cut off by pyrovalves of simple design, h regu-
lator and choke operating from electric drives are installed in the main
lines to control the thrust and ratio of fuel component flow rates during
flight of the engine. Supercharging assemblies, which are small gas genera-
tors in which gases are produced for supercharging the rocket fuel tanks,
are provided in the engine. Ihe assemblies for attaching the engine to the
rocket provide rotation of it in a plane parallel to the longittdinal axis
to control the flight direction of the rocket and its position. The engine
is similar to the Rl>-107 in height but has considerably smaller overall
dimensions in horizontal measurement. The specific weight of the iy5-253
engine is 7.7 kg/t.

Development of liquid rocket engines for the "Proton" space rocket was an
important achievement of space rocket technology. Such outstanding events
as near-earth orbits of the autanatic scientific stations "Proton" with
mass up to 17 t and the launch of automatic recoverable stations "Zond" for
orbiting the moon, delivery of lunokhods to the moon, launch of the auto-
matic interplanetary stations which took samples of lunar soil, i«*iich landed
on Mars and which became satellites of these celestial bodies, are related
to the use of this rocket.

Engines With Pressure Feed of Fuel

Besides liquid rocket engines with pump feed of fuel, different models of
which were described above, a few models of liquid rocket engines with
pressure feed are used in space rockets. Itiese engines, which operate on
high-boiling self-igniting fuels, are used primarily in the upj^er rocket
stages whose propulsion plants develop a thrust of several tons and may be
fired repeatedly during flight.

Of greatest interest ar^ong the indicated engines is the AJlO-138 engine
desioned by the Aerojet Company, Two of these engines with a total thrust
of 7-2 t were installed on the last stage of one of the space rockets of
the "Titan-3" family and were designed for repeated firing for 6.5 hours of
orbital flight. The indic^'ted rocket stage, called "Transstage, " is used to
solve numerous problens, including simultaneous insertion of several payloads
into different orbits (up to eight sattllites were injected) .

23

Iteliable operation of the iyrlO-138 engine and of the propulsion plant as
a. whole was provided to a large extent by the type of fuel used aw3 the
selected scheiratic diagram of the engine. Nitrogen tetroxide and aerozine,
which as we already know provide design simplicity of liquid rocket engines
and which are easily stored without losses \mder space flight conditions,
are used as the fi»l. fhe fuel is delivered to the chajuber at helium pres-
sure. &e chamber operates at a pressure of 7 atm a«d is made without a
regenerative cooling channel. For this reason the pressure in the fuel
tanks hardly exceeds the chamber pressure and the wei^t of the fuel tanks
is sufficiently small.

Ute chauRber of the described liquid rocket engine consists of a removable
laixing head which contains several h^jndred openings for atomization of the
fuel and is made of aluminum alloy, and a housing designed for cooling by
ablation and radiation of heat into the surrotinding atmosphere. Mass is
removed from the inner surface of the chanber by the hot gas flow during
Sblation. The cambustion chamber and the initial section of the nozzle
are cooled in this wanner (the remaining part of the nozzle is cooled by
radiation) . Itiis part of the chamber is manufactured on a production
iwndrel. An ablating polymer material — refrasil fibers impregnated with
a rubberized phenol resin, with subsequent vulcanization of these fibers,
is initially applied to the mandrel. A layer of heat insulation ™ asbestos
felt impregnated with phenol resin, which is also vulcanized, is then applied
and finally an outside pcwer sheathing ■ — fiberglas impregnated with epoxide
resin, is applied to this layer after which the mandrel is removed. The
power sheathing is designed for a temperature of not nwre than 180*0.

The total operating time {operating life) of liquid rocket engines is
limited by the permissible ablation of the roiterial and comprises 500 s.
This operating time is provided by creation of a protective gas-liquid filsi
near the chamber fire wall by atomization of th« fuel through peripheral
openings of the mixing head. The output section of the nozzle, coded by
radiation, is a thin-walled shell manufactured fr':-m refractory allocs:
niobium and titanium. This shell is heated to 90G-1100*C during operation
of the engine.

"The engine is fired by fuel control and auxiliary solenoid valves . The
engine is installed in a gimbal suspension on the rocket stage and is de-
flected during flight by electrcwechanical drives.

T^e comparatively high 'Specific impulse (3,000 m/s) of the ftJlO-138 en-^ine
at low gas pressure in the contjustion chamJber was achieved by expanding the
gases in the jet nozzle to a very low pressure. Based on data on the working
principle of liquid rocket engines contained at the beginning of the pamphlet,
one may conclude that the considered engine has large overall dimensions and
weight. Actually, its height is 2.0 ra, diameter is 1.2 m and specific
weight is approximately 25 kg/t. Thus, the AJlO-138 engine, which develops
a thnist of less than 4 t, is comparable in overall dimensions to the
I.R91-AJ-9 engine, whose thrust is more than 45 tj the AJ10-13B is consi-
derably inferior to the RB-119 engine in welqht ratio.

24

The Main Liquid Rocket Engines of Spacecraft

Ihe conditions under vrfiich these vehicles are fo'ind — weightlessness, deep
vacuum, space cold and the thermal radiation of the sun and planets — have
a significant effect on operation of space propulsion plants with liquid
rocket engines and largely determine their design and characteristics.

For example under weightlessness the fuel in the tanks is ramdofnly mixed
(if appropriate measures are not taken) with gas bubbles used to pressurize
the tanks. If gas enters the engine when it is being fired, the engine
operating process is disrupted with probable failure of its main assemblies.
To prevent this, special components which separate the gas and liquid are
provided in the fuel tanks of spacecraft engines: elastic bags, elastic
membranes, bellows and so on. the gas and liquid fuel may also be separated
by special design of the intake fuel devices (reticular or capillary) and
by creation of artificial gravity by auxiliary engines* when the engines
are fired.

Ihe deep vacuum under which liquid rocket engines of spacecraft operate is

the reason that the surfaces of the engine components leave gas molecules

adsorbed on the surfaces and also particles of structural materials, coatings
and lubricants. Variation of the frictional properties of the surfaces
caused by these processes may cause spontaneous weldix a of the contacting

metal parts. Subsequent settling of the removed metal particles on the
components of the engine electrical equipment may lead to short circuiting.

All these undesirable phenomena place specific limitations on selection of

engine materials, enqine configuration and its arrangement on the spacecraft.

Moreover, the negative effect of cosmic radiation on the materials and
working products used should be taken into account in some cases when de-
signing space propulsion plants with liquid rocket engines.

The outlined concepts are related to all liquid rocket engines operating in

space. The characteristics of specific models of these engines are deter- „
mined by the problems for solution of which they are designed.

Characteristics of the Main Liquid Rocket Engines of Spacecraft

liquid rocket engines which provide significant variation of spacecraft
velocities, usually related tc transferring the vehicles from an artificial
satellite orbit of the planet to a flight trajectory toward another planet
or into space, transfer of vehicles from flight trajectory to artificial
satellite orbit, landing on a planet and takeoff from it, transfer from an
intermediate to a different orbit and so on are discussed in this chapter, *"■'

*The problem of separating the fuel and gas must also be solved for liquid
rocket engines of the upper stages both during the first and during subse-
quent firings of the engines.

♦♦Operations which require multiple firing of propulsion plan*;s are fre-
quently carried out by WTT.

25

;'!,,; il.ClBlUlY Or' ....
UiGLNAL PAGE IS PWK

significant variation of spacecraft velocity occurs in a nui^er of cases
daring BsaneuverinQ operations an4 correction of flight trajecteary.

according to the resolwed problems, the main spacecraft engines may be
booster, retroflre, correcting, takeoff, landing airf so on. One eRginet
may also fulfill several fmsctions, iisciadJng auxiliary ftmctions.

Individual models of the c«isidered liquid rodtet engines differ considerably
between <»ach other in design and characteristics: they are laade both with
pim^ and with pressure feed of fue'' . the thrust of these engines enaaei^asses
a range fr<» tens of kilograexs to sevBral tons and may be both recelated and
tmregulated; they ©ay operate coatinuoiisly for tenths of a secoid aM
several thousand seconds; the nt»b€r of firincjs of these engines varies
from <m% to several tens. Faultless operation of all these engines in space
is provided to a large degree by the ose of high-boilir.g self- igniting
fcipropellant or of high-boilinq imnopropellant in thest.

Liquid rocket engines with pump feed of biprcyellant having regenerate irely
cooled Chalmers are used extensively in Soviet spacecraft. Mgines only
with pressure delivery of fuel, whose chaafisers are freg-uently cooled by
ablation or radiation and also by creatior* of a protectix-e gas-liqaid film,
are ysed in foreign spacecraft.

Soiw BOdeis of Soviet and Aiserican engines are presented below •«hich give
an idea aix;ut the main spacecraft engines. Ttse Soviet engines were deve-
loped under the leadership of A. M, Isayev,

The Engine of the First Spacecraft

This engine was a single-chasifcer liquid rocket engine with p«ap feed of
high~boiling self-igniting fuel, consisting of a nitrous oxide oxidizer
and aiaine-based f\jel. The engine developed a thrust of 1,614 kg and
specific impulse of 2,610 ro/s at a cos&ustion cha!Rber pressure cf approxi-
roately 57 atrr. The engine was part of the TDlr-1 retrofire rocket, which
developed a thrust to transfer the spacecraft froni artificial earth satellite
orbit to a reentry trajectory. The engine was located in the center of the
toroidal fuel tank pod. This configuration of the proptil;Hor. plant provided
coT^pactness and low weight of the design-

Fuel feed to the engine was provided during firing in orbital flight by
elastic iseinbranes installed in the tanks and \fi ,:•. separate the fuel and
pressurization gas (nitrogen) . The ccnditions for nonnal ignition of the
fuel in the combustion chamber were c-^atcd by insulating the coutbustion
chaiaber cavity from the surrounding mcditEir! by a thin muffler welded into
the jet nozzle in the region of ninimuR cross-section. This muffler ejected
the gases forced in ^^he coiJibustion cha5<4;er from the beginning of engine
rpe ration,

Itie TOU-1 was used on the first Soviet .spacecraft "Vostok," piloted by Yu. A.
Gagarin and g. £. Tltov, and in sc^.ewhat -modified tortr, o-\ ' e siibsequent
spacecraft of this tvpe and also on the '"^.^^khod" sp^'.cocraf v .

26

Figure 13. Ilie "V<»tok" Spacecraft With Last Rocket Stage: 1 —
recovery capsule-, 2 — retrofire rocket; 3 — • last
stage of "Vostok" roacet; 4 — liquid rocket stage

The TDO-1 joined to the -Vostok" spacecraft is shown in Fiqme 13.

The Engines of the Aatoaatic Interplanetary Stations "luna," "Venera" aad
-Mars-

fhe success ftal flit^.ts of Soviet aistomatic interplanetary stations to inves-
tigate the planets cf the solar syste« were provided by propulsion plants
*dth liquid rocket engines.

launches of aatanatic interplanetary stations to the soon were begun in 1959,
The problem of reaching the soon was solved initially. In this regard the
flight of the first three mm of the "Lima" type occurred on a ballistic
trajectory and they were not equipped with their owi engines. Solution of
more ccaiplex probless in investigation of the aoon, which science faced,
required development of special propulsion plants for automatic interplane-
tary stations and launch of vehicles frons earth parking orbits by firing
the engine of the last stage of the space rocket.

27

Hie Wis -wfeich made the world's first C1966} soft landing on the moon and
which went into circumluimr orbit were equipped with the KDTO~1 correcting-
retrofire propulsion plant. It included a single-chamber liquid rocket
engine with paoap feed of the same fuel as used in the engine of the "¥ostok"
spacecraft, "ate liquid rocket engine develc^^ed a thrust of 4,640 kg and
specific icfjulse of 2,720 a/s at a c«»bustion chaMser pressure of approxi-
aately 64 ata.

Fuel delivery to the liquid rocket engine without gas inclusions during
starting was provided by reticular separators in the tanks which utilize
the prqperty of the surface tension of liquid in the cells of a find metal
mesh. The enq."'' was fired twice during flight. The first firing provided
correction of the flight trajectory of the ms toward the iftoon and the
second provided braking to transfer the ftMS into an artificial lunar satel-
lite or to reduce the approach velocity of the JMCS to the moon to a safe
value .

The given fli^t direction and the specific position of the AMS in space
were provided by steering nozzles into which was fed the spent turbine gas,
these same nozzles provided (with a nonoperating chamber) the final approach
velocity of the ftMS to the lunar surface.

Later versions of the "Luna" type AMS delivered lunar soil specimens to the
earth and automatic self-propelled vehicles ''Luijokhod-l" and "Lunokhod-2"
to the lunar surface. Successful accocsplishment of these couplex problems
was provided by develops«nt of a new standardized attitude braking rocket
and, »»reover, a liquid rocket engine for the landing stage of the AMS
("Luna- 16" and "Luna-20'"}. The propulsion plant operated on high-boiling
self-igniting fuels containing asyufnetrical dirrethyl hydrazine and consisted
of two autoncaous pods (the main and a low-thrust pod) . The aain pod con-
sists of a single-chamber liquid rocket engine with punp feed of fuel and
was designed to operate in three different thrust modes (from 1,930 to
750 kg) and multiple firing (up to 11 times) . "flhe low-thrust pod contains
a two-chasjber liquid rocket engine for one-time firing with pressure deli-
very of fuel (by helium) ; it is also designed to operate in three different
thrust modes ^'rom 210 to 350 kg) . The specific im| alse for the liquid
rocket engine of the main pod is 3,080 m/s and the total operating time is
650 s; these parameters are equal to 2,490 ra/s and 30 s, respectively, for
the liquid rocket engines of the lo^'- thrust pod.

The propulsion plant is all-welded and the liquid rocket engine chamber of
the main pod is the load-bearing member to which all the other components
are attached.

The engine of the takeoff stage of the RMS "Luna" is a liquid rocket engine
with pump feed of the same fuel on which the low-thrust engine pod operates.
The engine develops a thrust of 1,920 kg at combustion chamber pressure of
approximately 94 atm. The launch of che takeoff stage of thf RKS "Luna-16"
is shown on the back cover.

28

•Sie leading role in the stn&f of ?ertus belongs to Soviet science. Launches
of AMS to this planet were begun in 1961 in the USSR, ftie fli#it schanes
initially provided entry of MfS into the ateesphere of 1?enus at escape ¥elo-
city, with separation of the descent capsules, braked by aerodynaasic drag,
fr<» the IMS. T3te flight trajectory of the W£ to ¥enus was provided by
a vernier rocket with liqttid rocket engine. Ifiis liquid rocket engine
CPigure 14) was one with pressure feed of fuel operating on a nitric acid
oxidizer and asfHi»et.rical dinethyl hydrazine, ffee engine was installed in
a gimbal suspension; it developed a thrust of 200 kg and specific impulse
of 2,670 m/s at chafer pressure of 12 at».

Figure 14. Vernier Engine of First M4S "Venera"

The described engine provided flight of "Venera" stations whose descent
capsules first made a smooth descent in the atinosphere of ¥enus (1967) and
which transmitted information frost the surface of this planet C1970) .
Similar liquid rocket engines were used in the space research progran using
AHS of the "Zond" series, the last variants of which c i re imnavi gated the
moon with subsequent return to earth. These same engines were used on the
"Holniya-1" coiwuni cat ions satellites and the first MIS launched toward
Mars.

The latest achievements of Soviet science and technology in investigation
of the planets of the solar system are related to development of more im-
proved propulsion plants with liquid rocket engines. Thus, vernier-braking

rockets with puirp feed of fuel, capable of reliaJble operation during the
many months of spaceflight in different thrust modes (from approxinsately
1,000 bo 1,900 kg) and during multiple firing {i^p to seven times), were
developed for the IKS "Mars," "Hie combustion chamber pressure of these
liquid rocket engines exceeds 95 &tm and the specific pulse reaches 3,090 m/s.

The "Soyuz" Spacecraft Engines

The "Soyu2" spacecraft is equipped with an approach-correcting propulsion
plant lAich includes two liquid rocket engines — a main and backup. The
main engine creates thrust to correct the orbit of the "Soyuz" spacecraft,
its maneuvers during approach to another spacecraft and to brake the space-
craft during descent froa artificial earth satellite orbit. Ihe backi;i)
engine operates in case of failure of the main engine and during isfjer-
missihle deviations in operation of the auxiliary engines (with -tetiich the
spacecraft is also equipped) ,

Both engines of the propulsion plant have puuf? feed ..r fuel Ca nitric acid
oxidizer and asyianetrical dimethyl hydrazine) with unregulated thrust of
approximately 415 kg. The specific pulse of the main engine is almost
2,770 m/s, the chamber pressure is 40 atm and the nozzle output pressure
is 0.04 atin. Unlike the two-chaittser backup engine, this engine is a single-
chamber type. The main engine is equipped with steering nozzles to idiich
the spent turbine gas is fed.

Ihis engine is the first liquid rocket engine with pump feed of fuel -sAich
permits reliable m.ultiple starts and operation both over a long period of
time (several hundred secou<fe) and in the short pulse mode (lasting tenths
of a second) .

The "»pollo" Spacecraft Engines

The "/i^ollo" spacecraft on which flights of teerican astronauts to the isKJon
were made Juring the period 1969-1972, is equipped with three main liquid
rocket engines according to the spacecraft diagram shown in Figure 15.

An ?kJlO-137 liquid rocket engine of the Aerojet Coipany is installed in the
service module of the spacecraft. It provides correction of the flight
trajectory to the moon, insertion of the spacecraft into circumlunar orbit,
transfer of the crew module from this orbit to a flight trajectory toward
earth (with return of the astronauts) and trajectory correction. The
liquid rocket engine develops a thrust of 9.3 t, its weight is 370 kg and
height is 3.«? m.

An LMDE liquid rocket engine of the Thomps on- Ramo-Woold ridge Company is
installed on yhr landing stage of the lunar module. The thrust of its
engine is regulated in the range of 4.5-0.45 t, if:: weight is 170 kg and
neight is 2.5 m. to RSlS liquid rocket engine of th« tocketdync Contpany
is installed on the takeoff stage of the module. The thrust of this engine
is 1.6 t, weight is 90 kg and height is 1.3 m. The engines of the lunar
module are shown in Figure 16,

30

Figure 15.

"Apollo" Spacecraft; 1 and 5 — main liquid rcscket
engines; 2 — service module; 3 — return vehicle
(crew module) ; 4 and 7 — aveciliari' liquid rocket
engines of jet control systeinsi 6 — lunar iRoduie

All three liquid rocket engines are based on common principles, the saae
as the Miopias engine described above ;in the section devoted to liquid
rocket space engines $ they are single-^chasber engines with pressure feed
Cby helium) by f lel Caitrogen tetroxide-aerozine) . The engines operate at
low chaaber pressure (7-8.4 atjaj and the chaiaber is designed for ablation
and radiation cooling and the use of a gas -liquid fuel flla. "Hie cojsfjara-
tively high specific impulse of the liquid rocket engines ( front 2,980 m/s
for the IM)E to 3,080 m for the AJ10--137J is achiewd by large jet nozzles.
The long operating life Cfrom 10 minutes for the RS18 to IS minutes for
the HIDE*) is provided by reducing the gas t«»perature in the combustion
chaffiber by selecting the corresponding ratio of fuel components.

The AJlO-137 and LHDE liquid rocket engines are installed in gitnbal suspen-
sions and say be deflected by electric ;drives. All three engines are fired
by oxidizer and propellant fuel values combined into single pods. These
valves and the other components of the liquid rocket engine control systejr.
are reserved to increase the engine reliability. If the propulsion plant
with AJlO-137 engine fails, the crew may be returned to earth by the IM3E
landing engine. This capability was realized successfully during the emer-
gency flight of an "Apollo" spacecraft to the moon in 1970 {"Apollo-13"} .

♦The liquid rocket engines operated less than the indicated time during
flights to the moon.

31

Figure 16. Arrangement of Main Liquid Itocket Engines of the Lunar
Module of the "Apollo" Spacecraft I the takeoff Stage of
the J*jdtile Is Shown on Topi : 1,7 and 9 — oxidizer
tanks? 2 — - liqudi rocket takeoff engine? 3 aM 10 —
assesfclies of the fyel tank pressurization system; 4
and 8 — heliias tanks for pressttrizing the fuel tanks,-
5, 6 and 11 — • fael tanks i 12 — liquid rocket landing
engine

ftie MS "Viking" Engines

l!ie American program for investigation of Mars including landing of two MS
"Viking" consisting of orbital and descent capsules,* on the surface in
Jtily-September 1976.

fhe orbital vehicle is equipped with an RS2101C ligtiid rocket engine of the
Eocketdyne Ccwpany, designed to correct the interplanetary flight trajectory
of the AMS, to insert it into the orbit of an artificial Martian satellite
and to correct this orbit. Itie indicated engine is a modification of the

♦Both "Vikings" have already landed on t*e surface of Mars aM are trans-
mitting scientific information.

32

engine used for the same purposes in the ms "Mariner-^," -(Aich photo-
graphed alstost 80 percent of the Ifertian surface frcro the orbit of this
planet in 1971.

Ihe RS2101C is a single-chanber liquid rocket engine with pressure feed of
a high-boiling self-igniting f«el (nitrogen tetroxide-monaoaethyl hydrazine) .
fhe engine is designed for 24 firings and a total operating tine of almost
1 hour Cthe length of one firing is 0.4 seconds to 45 aiin) . "flhe engine
develops a thrust of approximately 140 kg and specific impulse of more than
2,850 ib/s at chai&er pressure of 8 atm. The. engine is i«staiied in a gimbal
suspension. Hie weight of the engine is 8.2 kg, height is 0.6 m and dia-
meter is 0.3 ».

A soft landing of the descent capsule of the "Viking" spacecraft is provided
by three MR-80ft liquid rocket engines of the Socket Research Coaiany, which
operate on a monopropellant -~ hydrazine. The indicated engine is shown in
Fig«re 17. Its thrust is regulated over a range from 270 to 40 kg vdlth
operating time ijp to 10 min. A large number of jet nozzles (18) Is provided
in the engine to reduce the degree of daraage and cont»}ination of Martian
soil by the gas jet of the working engine.

—I — ' — ! — I — ; — < •'. ' ' 1 — I — I — r-T-T

* 5 S ? i S 19

Figure 17. Liquid Rocket Engine of Descent Vehicle of the "Viking'^
Spacecraft (One Scale Division Corresponds to 1.5 cm)

33

The indicated fact is related to one of the sain prdblems of the "Viking"
program — investigation of possible life on Mars, which prest»B»s sampling
and analysis of Martian soil specimens. Ihis problem predeteriained selec-
tion of hydrazine monopropellant (rather than biprope Hants) for the liquid
rocket engines of the landing mc^ule of the spacecraft. "She fact is that
the jet stream of the hydrazine engine has comparatively low temperature,
contains no undesirable by-products and, finally, the landing module with
the hydrazine -charged propulsion plant may be easily sterilized prior to
launch by holding it at an increased temperature — approximately 130 "C
Cthe vehicle is sterilized to eliminate possible transport of earth micix»~
organisms to Mars) .

Liquid rocket engines operating on hydrazine are discussed in detail in the
next chapter.

Auxiliary Liquid Rocket Fr.>^ines

These engines are used for orientation, stabilization of position, space-
craft and rocket stage trajectory correction, to separate fuel and pressuri-
zation gas (settling of fuel) in the fuel tanks prior to firing of the main
engines and so on.

Ttie auxiliary engines, similar to the main liquid rocket engines of space-
craft, operate on high-boiling self-igniting fuel or on monopropellajnt,
but only with pressure feed of it. The considered engines are significantly
inferior on the whole to other liquid rocket space engines in thritst level.
For exaniple, the ncminal thrust of the liquid rocket engines used to con-
trol the service module of the "Apollo" spacecraft is approximately 45 kg,
whereas the thrust of the main engine exceeds 9 t. Auxiliary engines with
a thrust of several kilograms or less are related to so-called microrocket
engines.

TSie most specific among all the auxiliary engines are those us*'-' for orien-
tation and attitude control of spacecraft Cthe rocket stages) . Several of
these liquid rocket engines, oriented by different control channels (pitch,
yaw and rolling) and grouped in pods, usually form the jet control system
of the spacecraft together with the comiron fuel tanks. The liquid rocket
enoines of jet control systems m.ay operate both in the continuous thrust
mode (with fixed or regulated value) and in the pulsed mode in which the
engine is fired periodically for a specific time. The pulse repetition rate
and pulse elngth may be quite different,

A characteristic feature of liquid rocket engines used in jet control systems
is the possibility of a very large number of firings, which may reach several
hundred thousand.

All the indicated characteristics of auxiliary liquid rocket space engines
are easily in the exaip.ple of specific models of these engines which will now
be considered. These engines are divided for con^'ienience into two groups
according t.o the type of fuel used in them. In conclusion we shall also

34

briefly discuss liquid rocket engines which cosmonauts amy use for propul-
sion in open space.

Itonopropellant Liqvild locket Engines

Liquid rotdcet space engiines of this type operate on low-water hydrogen
peroMide {concentration of approximately 90 percent) or hydrazine, <teco»-
posed in the chaanber of the engine in the presence of a catalyst} the
formed hii*~teBperature gas creates thrust %ftien it flows from the nozzle.

A typical diagram of a propulsion plant with monopropellant liquid rocket
engine is shown in Figxire 18, fhe fuel is stored in an elastic bag located
in the tank and enters the chamber under gaseous nitrogen or heliuui pres-
svure. The fuel decoisposition catalyst is located in the chmber itself.
Fuel feed to the chamber is controlled by a high-speed electric solenoid

"mwm-"

Figure 18. schematic Diagram of Jet Control System With Monopro-
pellant Liquid Rocket Engine: 1 and 10 — ■ filling
val'TOs? 2 • — elastic bag,* 3 — fuel tank,- 4, 6 a«d
8 -- control valves; 5 — filter,- 7, 9, 13 and 15 --
drain-safety valves? 11 — high-pressure gas tank;
12 — pressure reducer; 14 — drain valve; 16 —
liquid rocket engine chasiber

Auxiliary liquid rocket engines operating on hydrogen peroxide were used
first. Development of these engines was relatively siwple, since hydrogen
peroxide had already been used prior to this as a fsonofuel in rocket propul-
sion plants and as a source for producing gas to drive the turbopump assem-
blies. Moreover, liquid rocket engines were tested, for example, in the
United States for the jet control systews of experiffiental rocket aircraft.

The weight and operating characteristics of ircinoprop€ lant engines are deter-
mined to a great extent by the characteristics of the fuel decomposition
catalyst. Potassium and silver permanganate were used as catalysts in
auxiliary liquid rocket space engines.: In the first casB the catalyst

packet arranged in the engine chamber

Is a mixture of carrier substance

granules itifsregnated with potassium permanganate (Figure 191. In the second

35

case the catalifst packet is aade in the fosrm of a silver-coated wire frame
tisaally aaniifactured from stainless steel. Liqaid rodket engines witti
this type of catalyst packet are asre iifjroved.

kit fit*

Figure 19, Liquid Rocket Engine Chaiiser Operating on Hydrogen

Peroxide: 1 — monopropellaftt feed line? 2 — atomizer
head} 3 — catalyst packet; 4 -— jet nozzle

The fas tejaperatiu-^ in the decomposition chairJser is approxiiately 800*C and
the specific impulse is approximately 1,400-1,600 m/s for liquid rocket
engines operating on 90 percent hydrogen peroxide, Hi^er values of the
Indicated -w^alues correspond to a fuel with greater hydrogen peroxide con-
centration. It should be noted in this regard that the req«ir«Ments on
selection of the structural materials and the cleanliness of the working
components coming into contact with the fuel increase as the hydrogen
peroxide concentration increases to avoid spontaneous decoroposition of it.
The relatively low tenperature of the gases formed upon decomposition of
hydrogen peroxide permits irianufacture of the chamber for liquid rocket
engines operating on this fuel from ordinary stainless steel (by designing
them for radiant cooling) .

These engines (designed by the Bell Company) ^irere used, for example, in
1962-1963 on the "Mercury" spacecraft and have been tised since 1963 on the
"Centaur" rocket stage, A total of 18 liquid rocket engines joined into
two independent jet control systems — automatic and manual, were used on
the "Mercury" riacecraft. Three types of liquid rocket engines • — with
thrust of 0.45, 2,7 and li kg, operating in the pulsed mode, were used in
the automatic system. Liquid rocket engines of the latter two types, but
with regulated tl-irust, were used in the manual system, which was the backup
system.

Twelve auxiliary liquid rocket engines with thrust of 2.? kg each, joined
into four pods installed on the periphery of "he stage, are used in the

36

rocket stage of the "Centaur," They provide orientation of tiie stage,
settling of f«ei in the taxik and braking of the stagp upon separatic«i of
tlie spacecraft,

Hydrogen per<«ide was used esctensively in the auxiliary liquid rocket space
engines during the middle lS60's, after -which it gr»3ually began to be
replaced by hydrazine and liqiiid biprope Hants, which pennifc liquid rocket
engines to produce a hi^er specific infjulse,

The specific infsulse of liquid rocket engines was incresBied by approximately
40 percent upon conversion from hyd«>gen peroxide to hydrazine. Moreover,
imlike highly concentrated hydrogen peroxide, hydrazine is not subject to
spontaneous decc«i>as ition . The indicated advantages of hydrazine were
evaluated in 1958-1959, when developoent of liquid rocket engines for auto-
matic interplanetary stations was begun in the United States.

However, the nature of hydrazine deccaiposlticai was little sttdied, which
caiised difficulties in development of a catalyst which could actively decosi-
pose the hydrazine at room tenperature and tolerate multiple firings of
the engine. This cr,-alyst appeared in the United States in 1964. It was
developed by the Schell DevelopsKnt Company and was called "Schell-405."

The active substance of the catalyst which provides the coinposition is
iridiuir.. It is applied to porous granules of the carrier substance, which
is aluffiinum oxide.

Numerous models of liquid rocket space engines operating on hydrazine were
developed and found application with the appearance of the indicated cata-
lyst abroad, llie range of thrusts developed by these engines is from
approximately 20 g to 300 kg.

'She gas temperature in the decomposition chamber of hydrazine engines is
comparatively low (approximately lOOO^C) , #iich permits manufacture of the
engine chamber from refractory alloys, taking into account their radiant
cooling (as in the case of liquid rocket engines operating on hydrogen
peroxide). A typical hydrazine- fueled engine, developed by the Rocket
Research Conpany, one of the leading American companies in the field of
developing these engines, is shown in Figure 20. Hydrazine engines for
satellite control systems in which these engines are used extensively are
discussed below.

Hydrazine engines are used specifically ia commanications satellites opera-
ting in geosynchronous orbits. Itiese satellites are usually stabi^-ized by
rotation and several pairs of liquid rocket engines with a tbrust of approxi-
mately 2 kg, which provide given angular satellite rotational velocity
(usually 60-100 rpm) , turning of the rotational axis and also maintaining
and maneuvering of the satellite In the orbital plane, are used in them.

Liquid rocket engines may operate in both pulsed and in steady thrust modes,
ft typical pulsed mode Includes transr'lsE:'^n of a series of thrust pulses

37

Figure 20. Hydrazine-IPropelled Er.jine With Itiirust of 0.9 kg
fOplted States)

lasting 0.1 seconi with a patse of 0.9 second between th©ffl. Bie series nay
consist of several pulses and also of several hundred pulses. Itielir total
nurober coa^jrises sei«ral tens of idiousands, fhe total length of liqiaid
rocket engine operation in the steady thriist i»de reaches several hotirs.
Mquid rocket engines are designed to operate for several years.

For example, SEA16~6 liquid rocket engines of the Hamilton Standard Coffipany
with a thrust of 2.3 kg are used in the "Skyiiet-2" satellite; their weight
is 320 g, hei^t is 17 era and diaineter is 3 en (the fuel reserve for opera~
tion of these engines is only 23 kg) . M]RE~4A liquid rocket engines of the
'aioHiJsoii~Raj»o~WooIdridfe Company with thrust of 1,6 kg were used on the
"Intelsat- 3" sateliitei their weight is : 245 q, height is 11 en and diaMter
is 2.5 c». "Sue fuel tanks on this satellite are Installed so that the cen-
trifugal force developed during rotation provides reliable separation of the
fuel and pressurization gas, which is located in the fuel tanks thessselves.

Kfdragine engines with the lowest thrust are used in satellite orientation
systems stabiiiued in three asces. Itit thrust of these engines comprises
less than 50 g? they are designed for an operating life up to 450,000 working
pulses and may be operated up to 7 years. Pour of th -se engines together
with 16 hydrazine engines having a thnast of approximately 500 g are used,
for exawple, in the "FleetsatcoRt" conauni cations satellite? the hydrazine
is stored in fuel tanks together with the pressurization gas (hydrogen) and
is separated from the latter by an elastomer diaphragm. Electric heating
• • the entire satellite propulsion plant i^ provided to maintain the cata-
•.fst packet of the li<pid rocket engine at approximately 320*c. ifie need
for heating is related to the fact that a large number of starts of hydra-
zine engines with a cold catalyst leads; to breakdown and loss of quality.
It should be noted In completing the surroy of hydrazine engines that, de-
spite the intensive development of these engines # the process of tiieir
development is primarily enfsirical in mature. Each type of hydrazine engine
design works well only in a strictly determined mode and it is impossible
to predict the extent to i*ich it will satisfy new requirements.

38

Bipropellant liquid Rocket Engines

Ihe amin advantages of propulsion plants with inoriopro|)£liant liqtiid rocket
engines is their comparative siRf)licity and consequently the high design
reliability. Tnese propulsion plants also have weight advantages over
other jet systems within a specific range of values of total thrust inpulse
(i.e., multiplication of thrust by the total operating time). Bipropellant
liquid rocket engines are advantageous as the required total thrust impulse
increases with regard to their higher specific impulse.* Moreover, no
catalyst, which decensines to a great extent the dynamic characteristics
of the working process and which limits the operating life of the engine,
is required to operace these liquid rodcet engines.

Auxiliary bipropellant liquid rocket engines have now become as coiimon in
cosuonautics as monopropellant engines. Their developswnt was accelerated
to a large extent by tnanned space flight programs. Specifically, they
played an infjortant role in the "Apollo" program. A total of 52 auxiliary
liquid rocket space engines of four types with thrust from 34 to 68 kg was
used in the "Saturn- 5- Apollo" space rocket system. They operate at all
stages of space flight, beginning with launch of the "Apollo" spacecraft
to the ffioon and ending with return of i:^*> astronauts to earth.

The R~4d liquid rocket engines of the Markward Company, installed on the
body of the spacecraft service module, are shown in Figure 21. Ihis
module has four self-contained controlled subsystems with their own fuel
tanks and a compressed gas tank for fuel feed. The auxiliary liquid rocket
engines are grouped in four pods. Each pod weighs 18 kg and consists of
four R-4D engines arranged in a cross shape.

The indicated engine pods are part of the corresponding control subsystems
which are almost identical and form in aggregate a system which provides
orientation and stabiliza' ion of the service module and spacecraft as a
vrtiole, maneuvers of the module and other operations. Operation of the
subsystems is matched, which is required for uniform delivery of fuel.
Reliable spacecraft control is provided even if individual components of
the system fail, including failure of the subsystems.

The R-4D engines were also used in the lunar module control system of the
"llpollo" spacecraft. This syrtem also contains four liquid rocket engine
pods, but the engines are distributed between twc self-contained subsystemiS
of eight engines each. Ihe auxiliary engines may be switched to feeding
frwn the fuel tanks of the main propulsion plant if required.

The lunar module jet control system is installed on its takeoff stage (see
Figures 15 and 16) and provides separation of the lunar module ftom the
service module, orientation, stabilization and ma.neuverlng <ind also other

*Compressed~gas jet systems may be used at small values of total thrust
impulse.

39

.„, PAGE m POOH

Figure 21, Arrangement of the Auxiliary' Engines of the "iftpollo"
Spacecraft Service Module Jet Control Systec; 1 —
crew module; 2 — service ixjdule; 3 — auxiliary
liquid rocket engines

operations. The engines of the control syster. also provide a specific in-
crease in the speed of the takeoff stabe in the event of preirature shutdown
of the nain engine,

ItiQ P,-4D engine operates on aerozine-nitrogen tetroxide or moncr.cthyl hydra-
zine- nitrogen tetroxide fuel and develops 3 thrust of 45 )cg at a chamber
pressure of approxirately 7 atjp. Th« engine weighs 2.3 kg, its height is
34 cir, and its diameter is 16.5 cr,. The engine consists of a charier snd
solenoid oxidizer and fuel valves installed on it. There is a total of 70
ccr|:;onents in the engine structure.

The output section of the ligixid rocket engine jet nozzle is ranufactured
cf stainless steel and the reciaininc part of the chaniber housing is tnanu-
factured froir ixjlybdenuc", alloy and is covered on the inside with an anti-
cxidizing ceraroic composition fwclybdenur disilicide). The charsber is de-
signed for radiant coding in cctrbination with cooling by protective filr.
of fuel ani operates at g-'.s temperature of IISO^C, which is ccnsiderah-ly
below the temperature correspon-linc to production of raxiptm specific ir-
pulsc. The reduced temperature, v?hich provides charrber reliability, is
achieved by selcctino a snecific ratio of fuel and oxidizer flow rates.

T^le R-4D engines ray ftxctlcn in tcth the continuous and pulse thrust redes.
T- total of 40<^ of these engines, the total operating tir^r of which was

4^;

9 hours and the nurfjer of firings of i*ich was approxiiaately 750,000, was
used in the "Apollo" program. The total operating life exceeded 7 minutes,
the pulse repetition rate reached 30 per second with pulse length of 0.01
secoM and the noafcer of firings was 35,000 for individiial BKslels of the
engines Cat different flight stages}.

Strict evaluation of the utilization efficiency of the fuel consumed by the
engine during a single operating cycle is required with such a high nucfcer
of engine firings. Uneconcmicai fuel utilization during firing and shutting
off the engine (i.e., during transient cf>erating laodes) nay lead to a sig-
nificant decrease of its specific i^ulse and consequently to the need to
increase the fuel reserves on boaird the spacecraft and in the final analysis
to a decrease in the pay load.

TSje nature of firing and shutting down the engine is determined to a great
extent by its design. «ie» designing liquid rocket engines operating in the
pulsed aode, one must atte«f)t to a.ckieve a iTa.xi»t» control valve response
rate (miniRum tine) and mniroun voltnpe of spaces between valves and the
operating zone of the ccdb-ostion charJber (this is also desirable for other
reasons) .

fhB high speed of the R-4D engine control val-ws (their response tii« is
less than 0.01 s) is provided by the presence of only two wsving parts Ca
rod and spring) of small mass in each of thein. The miniowaK volume of the
mentioned spaces is achieved by installing the valves directly in the mixing
head (so that it is at the same tiiae part of the valve design) and in the
adopted chariser cooling scheme.

However, uneconomical fuel utilization during starting and shutting down the
engine could not be coi!f>letely eliminated (this is impossible with regard
to the nature of the cyerating process itself) . The specific irf^ulse of the
R-4D when operating in the continuous thrust mode is 2,750 k>./s and that when
operating in the pulsed aode is 1,360 m/s; the specific impulse during the
total operating tiiiie (average) is 2,160 a/s.

Microcracks, caused by the nature of its loading, usually appear in the
chamber design v«hen the engine is functioning in the pulsed isxxie. Develop-
ment of these K^iciocratcks limits the engine operating life. When selecting
Eiolybdenum disiliclde as the coating for the fi-4D chamber wall, its capa-
bility to "heal" the microcracks was ta3cen into account. This effect is
achieved by formation of silicon dioxide in the crack zone, which protects
the design material of the chamber wall against oxidation.

The R-4D is an example of a unversal liquid rocket engine. It was used
not only in the "Apollo" spacecraft but also in the "Lunar Orbiter" space-
craft (one of the ftmerican lunar programs) and also during flights of the
"Skylab" program. The P~4D engines installed in gimbal suspensions in the
"Lunar Orbiter" perfonned functions inherent to both auxiliary and inain
engines ,

41

tod now a few words about the SE~8 liquid rocket engine of the Rocketdyne
CwRpany. Twelve of these engines are used on the "J^ollo" spacecraft in
the crew modtile jet control syste». The engine operates on iwncarethyl
hydra2infr-'nitrogen tetroxide fxMl and develops a thrust of 42 kg. Hie
chanfcer is designed for ablation cooling in contfsination with fuel-film
cooling and operates at approximately the saxm teEf>erature as the R-4D
chastser. A refractory lining on a graj^ite base is installed in the mini-
mum cross-section of the chasiser. The maximum specific iufjulse of the
SE-8 is 2,650 m/s.

The E-4D aad SE-8 engines which we discussed are related to the largest
auxiliary engines operating on bipropellant. AB»ng the smallest engines
«ay be named those of the West German Messerschmitt-Belkoff-Blum CoHf»any,
tised in the orientation system of the "Syitphonia" coiwiunications satellite
(desired to operate for 5 years) . This engine weighing 160 g operates on
the same fuel as the SE-8 and develops a thrust of 1 kg at maximum specific
impulse of 2,870 nt/s.

Hquid Rocket Engines for Astronaut Propulsion

Investigations to develop liquid rocket engines to propel astronauts in open
space are being intensively conducted by Auierican coBapanifs. The MHO back-
pack propulsion unit, wAiich operates on hydrogen peroxide, was successfully
tested in 1966 during flight of one of the "Gemini" spacecraft. It contained
12 liquid rocket engines of the Walter Kid Company with a thrust of 1 kg each.

Arming the latest developH«nts shoiild be mentioned the HHHMU propulsion unit
of the Rocket Research Company, which operates or. a hydrazine-water mixture.
Because of dilution of the fuel by water, the temperature of the c^ses
ejected from the engine is only 260*C and the unit is safe for the astronaut.
It consists of two liquid rocket engines having a thrust of 0.45 kg and two
having a thrust of 0.9 kg. The entire unit is a small device with a compact
fuel tank which the astronaut holds in his hand. The fuel tank is easily
replaced with a new one; this operation may be performed outside the space-
craft.

Prospects for Development of Liquid Rocket Space Engines

an important role is being allocated in plans for further conquest of space
by liquid rocket engines. Powerful liquid rocket engines designed for
econc»tic utilization of hiahly efficient fuels are as before at the center
of attention of specialists. Proof of this is the SSME oxygen-hydrogen
enoine with thrust of more than 200 t, designed to transport th*^ "Space
Shuttle" spacecraft and developed by the Rocketdyne Company. Thi;- engine,
unlike the previous ox:;/gen-hydrogen engines, should function over the entire
spacecraft acceleration leg: from start to near-earth orbit insertion. The
chamJaer pressure of the SSMF engine exceeds 200 atm. The engine is now in
the bench development stage.

Development of cryogenic technolooy along with advances in the field of
heat-insulatinc materials will soon riiCde development of main and auxiliary
engines, operating on cryogenic fuels and developing a high specific im[ulse,

42

feasible for spacecraft. A mixture of liquid hydrogen with solid, so-called
slush-like hydrogen is of great interest. Rocket fuel density increases
%*ten liquid hydrogen is replaced by solid hydrogen and the hydrogen losses
to evaporation axe reduced several times,

A lot of attention is also being devoted to developiwnt of liquid rocke^
engines designed for use of new wore efficient fuels. Significant progress
has new been achieved in development of liquid rocket engines i^ith a thrust
of several tons, which operate on fwels containing fl«»rine — the strongest
of known oxidizers. Fluorine-hydrogen fuel exceeds oxygen-hydrogen fuel by
approximately 5 percent in specific impulse and doiJsle in density. The use
of fluorine fuels in which the oxidizer is fluorine, fluorine monoxide or
iKichanical mixtures of fluorine and oxygen (fluxes) and the fuel is hydra--
zine, aj3SK>nia, diborane or light hydrocarbons {for exanple, methane), may
produce a significant effect in sojse cases. These fuels provide a specific
impulse approximately 10 percent less than oxygei.-hydrogen fuel, but their
density is just as high as oxygen -kerosene fuel; moreover, they store better
under space flight ccmditions than hydrogen-containing fuels.

However, the development of fluorine fuels requires solution of nianerous
problems related to the chemical nature of fluorine, fhis product is excep-
tionally aggressive. It reacts with almost all organic and inorganic sub-
stai ces with release of hidi ajnounts of heat, which frequently causes com-
bustion. Under specific conditions fluorine even reacts with inert gases.
Kany metals react with fluorine even at room temperature and with insignifi-
cant heating (fluorine burns all metals with strong hea'iing in the atmo-
sphere). Asbestos, glass, sand and concrete bum in a fluorine jet; a fire
begun as a result of fluorine action is very difficult to extinguish.
Fluorine oxidizers and combustion by-products of fluorine fuels are unfor-
tunately related to the most toxic products dangerous to man and the environ-
■st.

The chetrdcal aggressiveness of fluorine considerably restricts the nuirber of
possible construction materials for liquid rocket engines designed to use
fluorine fuels. The inner surfaces of metal struct'iral components coming
into contact with fluorine should be subjected to passivation operation,
which includes treatment of the surfaces with gaseous fluorine to apply a
protective fluoride film. Ihe inner surfaces of the components should have
no pores, miarocracks, burrs and other defects, since they may cause ignition
of the structure. The use of fluorine fuels in liquid rocket engines com-
plicates the problem of developing a reliably cooled chamber since extremely
high temperatures are developed during combustion of these fuels (for example,
it is TOO^C higher for fluorine-ajroonia fuel than for oxygen-kerosene fuel
and reaches 4100'»C).

At the same time, despite the difficulty of developing fluorine fuels, the
advances achieved in this field during the past decade provide the basis to
assume that ^'luorine liquid rocket engines will find application during the
next few years in the upper stages of rockets designed to transfer automatic

43

spacecraft to other ojcfeits and to boost them toward the planets. Fluorine
liquid rocket engines are regarded as promising engines for spacecraft as
well which will be capable of itaking long flints to the planets.

Ozone, which is also a stronger oxidizer than oxygen, is being considered
among the fjossible fuel coi^ponents for future liquid rocket engines. Ozone
in combination with oxygen provides a theoretically higher specific impulse
than the fluorine-hydrogen fuel pair. However, ozone is an extretaely explo-
si^re product which tends strongly toward spontaneous detaiation and, conse-
quently, the problen of production and use of ozone in large quantities aiust
still be resolved.

Isfsroving the characteristics of bipropellant liquid fuels may also be
achieved by t-dding lightweight i»tals to them ms a third coirponent. Of
greatest interest ai»ng these inetai-cocitaining fuels are fluorine-hydrogen-
lithiuffi and ossygen-hydrogen-beryllium cosi>ositions , which essentially pro-
vide production of a specific iufjulse of approximately 5,000 m/s, close to
the maximum for existing molecular fuels. Such a high specific inpulse may
be explained siisply by the large aK>unt of heat released during combustion
of metals in oxygen and fluorine and by the low HKslecuiar weight of oxygen
which receives the dissipated heat.

Numerous problems ^ich include development of appropriate ir^thods of fuel
production and storage, organization of fuel feed to the engine chamber and
provision of total fuel coitiustion in the chamber with subsequent efficient
dispersion of the combustion byproducts, must be solved to develop econcsni-
cal engines operating on metal-containing fuel. A significant disad%rantage
of the two mentioned metal-containing fuels is their low density, caused by
the high hydrogen content (they are four times lighter than oxygen-kerosene
fuel).

Along with development of new rocket fuels, a search is being conducted
for engineering principles which provide further improvement of liquid roc-
ket engines in economy, overall dimensions and mass. The possibilities of
the schemes and design solutions adopted in nodern liquid rocket engines are
limited in this regard. The fact is that an advantage in the specific im-
pulse and overall diuKnsions of liquid rocket engines, achieved by increasing
the chamber pressure, becomes even less discernible as pressure increases and
the difficulties of developing liquid rocket engines increase more and more.
A significant increase of chamber pressure above 200-250 atm has little
effect and is difficult to achieve.

Liquid rocket engines with external expansion nozzle (Figure 22) are of great
interest in this regard. These engines are indebted to their name by the
fact that the gas in them flows past the jet nozzle on the outside of the
cljamber rather than inside the nozzle as in ordinary liquid rocket engines.
An exterior expansion nozzle is a shaped body which constricts the gas flow
in direction, similar to a tapered or prism-shaixid dish with a bottom. The
combustion chamber in Figure 22 is in the form of a ring encompassing the
nozzle. All the other corponents of the engine structure, including the

44

tiirbopamp assembly, are located inside the ncazle. "Hie spent turbine gas
is ejected to the outside throu#i openings in the bottwti of the nozzle.
«hen the rocket starts, the jet stream is initially pressed against the
nozzle by atmospheric pressure and then expands to the sides as the life
of the rocket increases. Since the gas flew past the nozzle eKpands to
pressure close to surrounding pressure, the nozzle operates constantly
during flight of the rocket in a n»de corresponding to the oaxiBniin specific
lispulse, s^ich is a significant advantage of an exterior expansion nozzle
ccwpared to those ordinarily used. Another advantage of an exterior expan-
sion nozzle is their significantly smaller overall diirensions (they are
three- four times shorter than ordinary nozzles), -jshich is explained by
their gas-dynamic characi-eristics. The use of exterior expansion nozzles
perffiits a significant increase of specific impulse and reduction of overall
dimensions of liquid rocket engines without resorting to an increase of
chaKiber presstire above 100 at©. Experiujental Ktodels of liquid rocket engines
with exterior expansion nozzle, designed for thrust of approxicmtely 10 to
100 t, have now been tested. It should be said that developinent of these
engines presents itany difficulties to designers and technologists.

Picrure 22. Mocfc-Up of Liquid Rocket Engine With Fxterior Expansion
Kozzle (United States)

The development of liquid rocket space engines is related to enormous mate-
rial expenditures. The cost of de\-elopino these engines reaches many hun-
dred million dollars and the cost of serially-produced models is frequently
expressed by seven- figure numhers. Nevertheless such expensive articles
together with other rocket and spacecraft components are used only once.
Investigations of projects for multi-use space transport systems, carried
out during the past few years, showed that their development is justified
economically only If they are used frequently. Invest i fiat ions of muiti-use
space svKteir.s recently entered the stage of developing the first models,
use of which is planned for the early 19B0*s. One of the key problems in

4^^

developing these systecis is to develop powerful efficient liquid rocket
engines designed for several tens of flights and for an operating life of
several hours with a small nxmber of adjustments between flights. Develop-
ment of a special system for technical diagnosis of the status of liquid
rocket engines becoaes necessary under these conditions, R new element
appears in the engine: a control unit with a sinall computer lAich controls
engine operation and issues instructions for energency shutdown if necessary.

"Hie first models of multi-use space systems, like existing spacecraft, will
be iRultistage. ScMne of their main coitf>onents as before are designed for
one-time use. An example is the Isiwrican "Space Shuttle" space transport
system. Multi~ise space systems which do not contain jettisonabie parts
and which confjletely justify their name,* will probably be developed by the
end of the centxiry.

The propulsion, plants of tJie systems indicated above and systeiss as a whole,
otiose operation brings a significant economic advantage, are in themselves
expensive facilities which embody the latest advances of science and techno-
logy. Along with this development of ordinary expendable rockets, irnperfect
in the engineering sense but not requiring large expenditures of funds and
time for deve Icpment , may be feasible for future space prograirs . M experi-
mental nK>del of an inexpensive liquid rocket engine for a rocket with thrtist
of 113 t, designed for pressure feed of fuel at a chamber pressure of 21 atm,
was tested on a test stand in the United States in 1968; this engine is
similar in design tn the landing engine of the "%!ollo" lunar module. 1!he
cost of the engine was a little m.ore than 20,000 dollars. Let us point out
for c<w|;arison that the serially-produced J-2 engine vrith thrust of 104 t
cost wsre than 1 million dollars.

Thus, the possibilities of developing liquid rocket space engines are far
from exhausted. It should be taken into account that we have considered
mainly only those prospects for development which are related to investiga-
tions already begun.

In conclusion let us say a few words about future liquid rocket space engines
in general, takinc into account the circumstance that other types of engines
capable of being used in cosmonautics exist and are being developed.

Conclusions

Liquid rocket space engines appeared .ibout 50 years after the idea of space
flight had been scientifically and technically advanced. Nurerous models of
liquid rocket space engines, diffcrina significantly between each ether in
external appearance, desioi and characteristics, were developed after launch
of the first artificial earth satellite. Along with engines which developed
a thrust of a fraction of a gram and which will fit in the r.alm, there are

*For more details see V. I. Levantovskiy 'p parr.nhlct "Transpcrtnyye
kosjnicheskiye sistemy" f?pace Transport Sy? terns 1 , Moscow, f-naniye, l^'^C-

4f-

engines 4*ith a thrust of 100 tons and height of several iseters ?i**5ich weigh
many tons, llie operating life of low-thrust liquid rocket engines reaches
several hours and the number of firings reaches many thousand, ^iiereas
powerful liquid rocket engines are usually fired once and operate less than
10 miniates.

Compared to M51T, also used extensively in cosmonautics, liquid rocket
engines have the advantage that they develop a higher specific iapulse and
can operate in their wjst diverse ttKSdes with multiple firings. The advan-
tages of RiyrT are determined by their design sijiplicity, ease of storage in
a charged state and the high density of solid fuel.

Liquid rocket engines and FDTT are related to chemical rocket engines, since
they develsqj thrust by using the potential chemical energy of the fuel, ihis
energy reserve restricts the specific iaipiilse of the engine to approximately
5 km/s, which is considerably less than orbital velocity. Irs this regard
enormous and complex rockets whose payioad is soiw fraction of the launch
5S, must be constructed for space flight.

A significant increase of the specific impulse of chemcal rocket engines
(up to 10-20 lar./s) may be expected only if hypothetical fuels are created
which contain free atoms and radicals or excited atoms and Kiolecules.
However, r«3l types of rocket engines with hiah specific impulse are known.
v?e are talking about nuclear and electric rocket engines.

The modern level of science and technology permits development of nuclear
rocket engines {YaRD) with a thrust of several tens of tons and specific
impulse of approximately 8 km/s. The working substance which develops
thrust in these engines is hydrogen, which is stored in the liquid state in
a tank. Upon passing through a nuclear reactor, the hydrogen is transformed
to a high-tenperature gas which is then accelerated in a jet nozzle. A
nuclear rocket engine may be used as the rrain engines of spacecraft and the
upper stages of space rockets. One of the most developed projects for using
nuclear rocket engines provides for expeditions to Fars.

Along with the working substance, electric energ>' used to heat or boost the
working substance, is employed in electric rocket engines {ERD5 to create
thrust. Because of the low thrust to weight ratio, EFD may be used only in
spacecraft. Mumerous models of electric rocket engines are now being used
as auxiliary space engines, they develop a thrust up to several kilograms
and specific impulse up to several tens of kilometers per second.

Investigations showed that electric rocket engines may also be used as the
rrain engines of spacecraft. With regard to the low acceleration imparted to
the spacecraft, these engines should operate continuously for a long time:
for exap-ple, more than 1 year when launchir.n a satellite to Jupiter.

Rational combination of nucloar and electric rocket engines with motlern
chemical engines will provide further progress in cosmonautics.

47

TABLE OF CONTENTS

General Data on the Working Principle of Liquid Rocket Space Engines . . 1

History of Development and Use of Liquid Rocket Engines 5

Main Liquid Rocket Space Engines .... 6

The engines of the first space rocket (the history of

development, design and parameters) 7

A brief survey of space rocket development 12

The F-1 liquid rocket engine 14

Oxygen-hydrogpn engines 16

Liquid rocket engines operating on hi^-boilinc fuel 19

The engines of the "Kosmos" and "Proton" space rockets 20

Engines with pressure feed of fuel 23

Main Liquid Rocket Engines of Spacecraft 25

■Hie characteristics of the main liquid rocket engines of spacecraft 25

The engine of the first spacecraft 26

The engines of the MAS "Luna," "Venera" and "Mars" 27

The engines of the "Soyuz" spacecraft 30

The engines of the "Apollo" spacecraft 30

The engines of the "Viking" spacecraft 32

Auxiliary Rocket Engines 34

Monopropellant rocket engines 35

Biprc^ellant rocket engines 39

Liquid rocket engines for astronaut propulsion 42

Prospects for Development of Liquid Itocket Space Engines 42

Conclusions ^^

4e